Ge Avio S.r.l.

Alessio Gargioli

Marco Garabello

Giuseppe Argentieri

Salvatore Garofalo

Paolo Altamura

Gian Mario Bragallini

Jan Biskupski

Fabrizio Bussi

Simone Castellani

Davide Lauria

Ottavia Pica

Simone Iurlaro

Enrica Gilardi

An aircraft engine component () may include a wall () comprising an aluminum alloy and/or a magnesium alloy, and a protective coating () covering the wall (). The protective coating () may include a prime layer (), a silicone elastomer layer (), and an abrasion resistant layer (). The prime layer () may at least partially cover a surface () of the wall (). The prime layer () may include a silane coupling agent and an organic titanate. The silicone elastomer layer () may at least partially cover the prime layer (). The silicone elastomer layer () may include one or more filler materials dispersed in a matrix of cross-linked silicone polymers. The abrasion resistant layer () may at least partially cover the silicone elastomer layer (). The abrasion resistant layer () may include a fiber-reinforced elastomeric material.

Protective coatings for aircraft engine components

A gearbox assembly for a gas turbine engine includes a planetary gear set comprising a sun gear, a planet gear, and a ring gear, the sun gear configured for connection to a first drive shaft of the gas turbine engine and the ring gear configured for connection to a second drive shaft of the gas turbine engine, and an electric machine assembly comprising an input and an electric machine. The electric machine comprising a rotor coupled to the input and a stator fixedly positioned within the gearbox assembly, the input of the electric machine assembly coupled to one of the sun gear, the ring gear, or the planet gear of the planetary gear set.

Hybrid transmission on propeller gearbox

An engine assembly defining an axial direction (A) and including a gearbox, an engine core including at least one rotor, and a flexible coupling shaft having a first end and a second end along the axial direction (A). The first end of the flexible coupling shaft is connected to the engine core and the second end of the flexible coupling shaft is connected to the gearbox. A damper system is positioned at the second end of the flexible coupling shaft. The damper system is configured to reduce vibrations to the flexible coupling shaft during operation of the engine assembly.

Damper system for an engine shaft

The present disclosure is directed to a turbine engine () defining an axial direction and a radial direction. The turbine engine includes a fan or propeller assembly () comprising a gearbox; an engine core () comprising one or more rotors, wherein at least one of the rotors defines an axially extended annular hub; and a flexible coupling shaft () defining a first end and a second end along the axial direction, wherein the first end is connected to the engine core and the second end is connected to the gearbox, and further wherein the flexible coupling shaft extends from the one or more rotors to the gearbox in the axial direction and inward of the hub in the radial direction.

Method and system of connecting a turbine engine gearbox to engine core

The present disclosure is directed to a shaft assembly () for a turbine engine (), wherein the turbine engine includes a fan or propeller assembly () and an engine core (), and further wherein the fan or propeller assembly includes a gearbox (), and wherein the engine core includes one or more rotors (). The shaft assembly () includes a flexible shaft () defining a first end () and a second end () along the axial direction, wherein the first end is connected to the engine core () and the second end is connected to the gearbox (), and wherein a plurality of splines () is defined at the second end () and coupled to a spline interface () at the gearbox (); and a coupling () extended at least partially in the radial direction and coupled to the engine core and the flexible shaft.

Flexible coupling shaft for turbine engine

The present disclosure is directed to a turbine engine () defining an axial direction, a radial direction, a circumferential direction, a first end () and a second end () opposite of the first end () along the axial direction. The turbine engine includes a propeller assembly () proximate to the first end including a plurality of blades () arranged in the circumferential direction disposed around an axial centerline (), and a feathering mechanism () including a hollow piston rod (). The feathering mechanism rotates the plurality of blades about a pitch axis () extended in the radial direction from the axial centerline. The turbine engine further includes a housing () proximate to the second end disposed in adjacent arrangement with the propeller assembly in the axial direction. The axial centerline is defined through the propeller assembly and the housing. The turbine engine further includes a beta tube assembly () extended through the hollow piston rod and at least partially through the housing in coaxial alignment with the axial centerline. The beta tube assembly defines an at least partially hollow walled pipe () extended along the axial direction. The beta tube assembly further defines a plurality of grooves () extended along the axial direction proximate to the housing. A first groove () extends at least partially in the circumferential direction and along the axial direction to at least partially define a helix () corresponding to a rotatable range of the plurality of blades about the pitch axis, and a second groove () extends in the axial direction.

Growing community of inventors

Turin, Italy

Alessio Gargioli