United Technologies Corporation

Raytheon Technologies Corporation

Pratt-whitney Canada Corp

Rtx Corporation

Sikorsky Aircraft Corporation

Gregory E Reinhardt

Enzo DiBenedetto

John C DiTomasso

Marc J Muldoon

Robert M Lewis

Todd A Davis

Gabriel L Suciu

William G Sheridan

Brian D Merry

Rajendra K Agrawal

Daniel W Major

Jeffrey Jay Lienau

Andrew Pope

Gabrielle Murphy

Karl D Blume

Gregory M Dolansky

Gregory M Savela

Joseph John Swiderski

Paul Thomas Rembish

Douglas Pollock

Jun Shi

Steven D Porter

Jorge I Farah

Anthony P Cherolis

Carl Brian Klinetob

Seth A Max

Kevin E Green

Barry William Spaulding

Jason W Mazanec

Donald Summers

Scott L Peterson

Real Barrette

Stephane Mongeau

Daniel L Gysling

Paul M Lutjen

Raymond C Surace

Paul R Hanrahan

Ioannis Alvanos

Daniel A Mosher

John E Holowczak

Thomas J Robertson, Jr

Joshua Daniel Winn

Nigel David Sawyers-Abbott

Steven J Feigleson

Christopher M Valva

Dmitriy A Romanov

Kevin Zacchera

Daniel S Rogers

Gary L Grogg

Daniel Benjamin

Thomas B Hyatt

Jeremy Drake

John H Mosley

John P Tirone

Brian C Lund

Taryn Narrow

Steven W Trinks

Jason Leroux

Brian P Huth

John P Lucashu

Enzo Dibenedetto

Steven L Conner

Steven G Lemieux

David J Yudichak

Collin P Fitzgerald

Leonard Kolshak

Réal Barrette

Stéphane Mongeau

Daniel A Szkrybalo

Nigel D Sawyers-abbott

Shi Jun

Rajendra Agrawal

Real Barette

An assembly is provided for an aircraft. This aircraft assembly includes a geartrain, a first bladed rotor, a second bladed rotor and a rotating structure. The geartrain includes a sun gear, a ring gear, a plurality of intermediate gears and a carrier. The ring gear circumscribes the sun gear and is rotatable about an axis. Each of the intermediate gears is between and meshed with the sun gear and the ring gear. Each of the intermediate gears is rotatably mounted to the carrier. The carrier is rotatable about the axis. The first bladed rotor is coupled to the ring gear. The second bladed rotor is coupled to the carrier. The rotating structure is coupled to the sun gear. The rotating structure includes a turbine rotor. The rotating structure is configured to drive rotation of the first bladed rotor and the second bladed rotor through the geartrain.

Selective power distribution for an aircraft propulsion system

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 first shaft supports a low pressure compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A first bearing supports the first shaft relative to the inlet case. A second bearing supports a second shaft relative to the intermediate case. A low pressure compressor hub is mounted to the first shaft. The low pressure compressor hub extends to the low pressure compressor section between the first bearing and the second bearing.

Gas turbine engine shaft bearing configuration

A machine has a first member; a second member; a third member; a bearing having an inner race mounted to the second member and an outer race mounted to the third member; a damper chamber between the first member and the third member; a fluid outlet in the first member to the damper chamber; a fluid supply flowpath to the fluid outlet; and an unvented chamber open to and locally above the fluid supply flowpath.

Squeeze film damper acoustic isolation

In one exemplary embodiment, a center body support for a gas turbine engine includes a fan exit guide vane mounting ring and an outer annular wall. An aft flange extends radially outward from the outer annular wall. An inner annular wall is radially spaced from and structurally interconnected to the outer annular wall by a circumferential array of struts. A brace interconnects the aft flange to a forward portion of the outer annular wall. The aft flange includes inner and outer mounting features spaced radially apart from one another. The inner and outer mounting feature respectively receive first and second sets of fasteners. The forward portion extends in a radially outward direction from the outer annular wall. The forward portion includes a forward mounting feature that receives a third set of fasteners. The outer and forward mounting features are secured to the fan exit guide vane mounting ring respectively by the first and third sets of fasteners. Forward, outer and inner mounting features define a triangular cross-sectional shape. The forward portion has a radial inner end opposite a radial outer end. The brace is joined at the radial outer end. The brace transfers an axial load from the aft flange to the radial outer end of the forward portion.

Gas turbine engine front center body architecture

A gas turbine includes a shaft, a turbine coupled with the shaft for rotation with the shaft, and a bearing coupled with the shaft to facilitate rotation of the shaft. A bearing nut is adjacent the bearing on the shaft. The turbine has a first load path and the bearing has a second load path. The bearing nut exerts a force on the bearing such that the first load path is not aligned with the second load path relative to a central axis of the gas turbine engine. A method of assembling a gas turbine engine is also disclosed.

Turbine bearing stack load bypass nut

A horizontal support tool for an engine build stand, the horizontal support tool includes a support tube along an axis and a tie shaft between a handle and a puck assembly, the puck assembly including a puck selectively extendable and retractable transverse to the axis in response to rotation of the handle. A method of horizontally assembling a portion of a gas turbine engine including mounting a first module to an engine build stand; installing a horizontal support tool into the first module, the horizontal support tool supported in a spherical bearing supported by the engine build stand; and installing a second module to the first module, the horizontal support tool operable to at least partially support second module.

Horizontal engine build stand

A bearing assembly for a gas turbine engine includes, among other things, a bearing housing extending along an axis to define a bearing compartment, a lubricant seal assembly adjacent to the bearing housing to bound the bearing compartment, an air seal assembly defining a vent cavity along the bearing housing, wherein a mixing cavity is defined between the lubricant seal assembly and the air seal assembly, and wherein the bearing housing defines an airflow supply passage, an airflow vent passage and a scupper passage having respective fluid ports at different circumferential positions relative to the axis, the fluid port of the airflow vent passage fluidly coupled to the vent cavity, and the fluid ports of the airflow supply and scupper passages fluidly coupled to the mixing cavity. A method of sealing for a gas turbine engine is also disclosed.

Bearing compartment oil auto-ignition mitigation

A gas turbine engine includes a fan that includes a plurality of fan blades rotatable about an axis. An on-wing portion includes a compressor section and a combustor that is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. A core flow path is arranged within a core nacelle. The fan is arranged upstream from the core flow path. A fan drive gear system module is coupled to the turbine section for rotating the fan about the axis. A connector assembly including first and second members respectively is secured to the on-wing portion and the fan drive gear system module. The first and second members are removably secured to one another by radially extending fasteners. The first members are connected to an on-wing portion of the gas turbine engine and the second members are connected to the fan drive gear system module. The fasteners are accessible through the bypass flow path.

Fan drive gear system module and inlet guide vane coupling mechanism

An assembly system for assembling a gas turbine engine including a support beam defining a horizontal plane for assembly of engine components along a horizontal axis substantially parallel with the support beam. A forward arm is supported on the support beam for supporting a forward portion of an engine component. An aft arm is supported on the support beam for supporting an aft portion of the engine component. An aft mounting ring is attachable to an aft end of an engine assembly. A forward mounting ring is attachable to a forward end of the engine assembly. A drive is mounted to the aft arm engageable to the aft mounting ring for rotating the engine assembly about the horizontal axis. A roller is supported on the forward arm and engageable to the forward mounting ring to support the first end of the engine assembly during rotation about the horizontal axis. A method is also disclosed.

Gas turbine engine assembly method and system

An engine mounting arrangement is provided for supporting a gas turbine engine. The mounting arrangement comprises a module adaptor mounted to an overhead or a ground based carrier for rotation about a substantially horizontal axis. The module adaptor extends horizontally from a base to an axially-opposed distal end having a first mounting interface configured for attachment to a module of the engine. The module adaptor holds the engine by the module to which it is attached and with the other engine modules projecting in a cantilever fashion from the supported module.

Gas turbine engine module adapter to a carrier

A horizontal support tool for an engine build stand, the horizontal support tool includes a support tube along an axis and a tie shaft between a handle and a puck assembly, said puck assembly including a puck selectively extendable and retractable transverse to the axis in response to rotation of the handle. A method of horizontally assembling a portion of a gas turbine engine including mounting a first module to an engine build stand; installing a horizontal support tool into the first module, the horizontal support tool supported in a spherical bearing supported by the engine build stand; and installing a second module to the first module, the horizontal support tool operable to at least partially support second module.

A sealing arrangement includes a turbine static structure with a seal ring having groove with a contact surface. Also included is a bearing compartment with a second contact surface, and a piston seal positioned between the turbine static structure and the bearing compartment. The piston seal includes sides configured to contact the contact surfaces and is positioned in the groove.

Bearing compartment seal

A turbofan engine structural guide vane is mounted to a forward bulkhead of a core engine case structure at an inner end and to a fan case at an outer end. A plurality of shear pins extend from the aft portion of the structural guide vane into a corresponding plurality of openings defined in the bulkhead for bearing circumferential loads.

Structural guide vane circumferential load bearing shear pin

A mid-turbine frame buffer system for a gas turbine engine includes a mid-turbine frame that supports a shaft by a bearing. An air compartment and a bearing compartment are arranged radially inward of the mid-turbine frame. The bearing compartment is arranged within the air compartment and includes first and second contact seals arranged on either side of the bearing. The air compartment includes multiple air seals. A high pressure compressor is fluidly connected to the air compartment and is configured to provide high pressure air to the air compartment. A method of providing pressurized air to a buffer system includes sealing a bearing compartment with contact seals, surrounding the bearing compartment with an air compartment, and supplying high pressure air to the air compartment.

Mid-turbine frame buffer system

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South Glastonbury, CT, United States of America

Gregory E Reinhardt