United Technologies Corporation

Other

Raytheon Technologies Corporation

James W Norris

Gabriel L Suciu

Craig A Nordeen

Brian D Merry

Wendell V Twelves

James D Hill

Bernard J Raver

Christopher Marc Dye

Ioannis Alvanos

Arthur M Salve, Jr

Stephen P Muron

Steven J Sirica

Gary D Roberge

Daniel Bernard Kupratis

William K Ackermann

Stephen Karl Kramer

Alfred Paul Matheny

Michael James Schmidt

Anton G Banks

Joseph W Bridges, Jr

Andrew P Berryann

Gary D Jones

James Chambers

Gary A Clawson

Brian Hoel

James E Jones

Gary M Stetson

Douglas A Thonebe

Wei-Hua Pan

Joseph F Dolan

David P Schofield

Caroline Westwater

Bradley C Sammann

Harvey L Klein

James Hurchala

Noel H Davis

A rotor for a gas turbine engine includes a plurality of blades which extend from a rotor disk at an interface, where the interface is defined along a spoke. A spool for a gas turbine engine includes the rotor disk, the plurality of blades with the interface defined along the spoke radially inboard of a blade platform, a rotor ring axially adjacent to the rotor disk, and a plurality of core gas path seals which extend from the rotor ring. Each of the plurality of core gas path seals extends from the rotor ring at a seal interface, with the seal interface defined along a spoke and the plurality of core gas path seals being axially adjacent to the blade platform.

Spoked rotor for a gas turbine engine

A non-metallic tailcone in a tip turbine engine includes a tapered wall structure disposed about a central axis. The non-metallic tailcone is fastened to a structural frame in the aft portion of the tip turbine engine. The tip turbine engine produces a first temperature gas stream from a first output source and a second temperature gas stream from a second output source. The second temperature gas stream is a lower temperature than the first temperature gas stream. The second temperature gas stream is discharged at an inner diameter of the tip turbine engine over an outer surface of the tailcone. Discharging the cooler second temperature gas stream at the inner diameter allows a non-metallic to be used to form the tailcone.

Tip turbine engine composite tailcone

A rotor for a gas turbine engine includes a plurality of blades which extend from a rotor disk, each of the plurality of blades extend from the rotor disk at an interface, the interface defined along a spoke.

A cooling system for cooling a fluid reaction apparatus of a gas turbine engine includes a vapor cooling subsystem and a film cooling subsystem. The vapor cooling subsystem has a vaporization section and a condenser section for cooling a portion of the fluid reaction apparatus. The condenser section is cooled by a fluid. The film cooling subsystem is configured for cooling a portion of the fluid reaction apparatus by discharging fluid out of openings defined in the fluid reaction apparatus. At least a portion of the fluid used to cool the condenser section of the vapor cooling subsystem is discharged out of the openings of the film cooling subsystem.

Hybrid vapor and film cooled turbine blade

A non-metallic tailcone () in a tip turbine engine includes a tapered wall structure disposed () about a central axis. The non-metallic tailcone is fastened to a structural frame () in the aft portion of the tip turbine engine. The tip turbine engine produces a first temperature gas stream from a first output source and a second temperature gas stream from a second output source. The second temperature gas stream is a lower temperature than the first temperature gas stream. The second temperature gas stream is discharged at an inner diameter of the tip turbine engine over an outer surface of the tailcone. Discharging the cooler second temperature gas stream at the inner diameter allows a non-metallic to be used to form the tailcone.

A gas turbine engine radial impeller includes first and second impeller portions that are secured to one another along a neutral bending plane of the radial impeller. A vapor cooling cavity is provided between the first and second impeller portions. The neutral bending plane is arranged in the vapor cooling cavity.

Gas turbine engine vapor cooled centrifugal impeller

A tip turbine engine provides increased efficiency while eliminating or reducing the number of axial compressor stages by moving the core airflow inlet aft of the fan. As a result, the core airflow entering the core airflow inlet is the fan exhaust, which is already at an increased pressure. A portion of the fan exhaust is guided radially inward, then axially forward and then radially outward through compressor chambers in the hollow fan blades for further, centrifugal compression.

Tip turbine engine with reverse core airflow

An integrally bladed rotor has an outer rim with a plurality of blades extending radially outwardly of the outer rim. A plurality of channels are formed radially inwardly of the outer rim. A discontinuity formed at a radially outer surface of the outer rim includes a first thin slot at a radially outer face of the outer rim with an enlarged seal holding area. A second thin slot is positioned radially inwardly of the seal holding. The first and second thin slots are thinner circumferentially than the enlarged seal holding area. A seal is inserted into the seal holding area. The seal does not extend into the first and second thin slots, nor into the channels.

Integrally bladed rotor with slotted outer rim

An integrally bladed rotor for a gas turbine engine includes at least one discontinuity formed in an outer face of an outer rim. The discontinuity reduces hoop stress in the outer rim.

A tip turbine engine assembly () includes a fan () rotatable in a fan plane () about an engine centerline (). The fan () includes a fan blade () that defines a core airflow passage therethrough. An engine case () surrounds and supports the fan and engine. A plurality of engine mounts () for mounting the engine to an aircraft are positioned circumferentially about the outside of the engine case in a single mounting plane () that is perpendicular to the engine centerline.

Tip turbine single plane mount

A turbine engine includes a plurality of variable fan inlet guide vanes. Where the turbine engine is a tip turbine engine, the variable fan inlet guide vanes permit the ability to control engine stability even though the fan-turbine rotor assembly is directly coupled to the axial compressor at a fixed rate. The fan inlet guide vanes may be actuated from an inner diameter of the fan inlet guide vanes.

Variable fan inlet guide vane assembly, turbine engine with such an assembly and corresponding controlling method

An inner diameter vane seal for a gas turbine engine comprises an annular, ring-like body having inner and outer diameter rims, forward and aft faces and an air passage. The outer diameter rim extends circumferentially for engaging inner diameter ends of stator vanes. The inner diameter rim extends circumferentially and is spaced radially from the inner outer diameter rim. The forward and aft faces extend radially between the outer diameter rim and the inner diameter rim. The air passage extends from the forward face to the aft face between the inner and outer diameter rims.

Rotating vane seal with cooling air passages

A tip turbine engine according to the present invention includes a plurality of independently variable inlet guide vanes for the fan and/or for the compressor. An actuator is operatively coupled to each of the flaps, such that each actuator can selectively vary the flap of its associated inlet guide vane. In one embodiment, the inlet guide vanes each include a pivotably mounted flap that is variable independently of the flaps of at least some of the other inlet guide vanes. In another embodiment, the inlet guide vanes each include at least one fluid outlet or nozzle directing pressurized air, as controlled by the associated actuator, to control inlet distortion.

Individual inlet guide vane control for tip turbine engine

A rotor for a gas turbine engine includes a plurality of blades which extend from a rotor disk and at least one spacer adjacent to the plurality of blades. A flow passage is defined between the rotor disk and the blades and spacer. A plurality of inlets are formed within the spacer to pump air into the flow passage.

Secondary flow arrangement for slotted rotor

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.

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Lebanon, CT, United States of America

James W Norris