Allison Advanced Development Company

Rolls-royce North American Technologies Inc.

Other

Gm Global Technolgoy Operations, Inc.

Allison Engine Company, Inc.

Brian Paul King

Robert Anthony Ress, Jr

Edward Claude Rice

Robert T Duge

Mark Stephen Henry

Craig E Heathco

Bruce A Ewing

Charles H Smale

Jason E Dundas

Preston Earl Light

An aircraft powerplant is disclosed that can be operated in at least three modes including as a gas turbine engine, as an engine having a ramburner, and as an engine having a forward compression combustor engine such as a ramjet and/or scramjet. An airflow valve is provided to direct air to a downstream portion of the aircraft engine and can be positioned as a function of the aircraft operating modes. The valve can be used to cocoon the gas turbine engine.

Turbine based combined cycle engine

One embodiment of the present invention is a gas turbine engine. Another embodiment is a gas turbine engine vane system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engine vanes. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Gas turbine engine vanes

A passive clearance control system for a gas turbine engine component. The system provides a blade shroud that is isolated from the component case and can move radially to maintain blade clearance. A plurality of vane sectors define a stator ring including a blade shroud. The stator ring and the control ring are coupled together and mounted within the component by a plurality of radial extending mounting members.

Compressor casing with passive tip clearance control and endwall ovalization control

A method for mechanically cutting a continuous circular shaped groove in a foil sheet for use in a metal matrix composite product. In one form the foil sheet is held adjacent a rotating machine tool by vacuum and a cutting tool is moved relative thereto to cut a spiral groove.

Method of manufacturing a metal matrix composite structure

The present invention contemplates an apparatus for forming a combustion mixture in a gas turbine engine. In one form, the apparatus includes a diffuser having at least two flowpath structures spaced apart to define a flowpath for directing fluid flow. At least one of the flowpath structures includes a flowpath surface and a plurality of edges disposed along a trailing end portion of the flowpath surface. The edges extend perpendicularly from the flowpath surface and are arranged generally parallel to the fluid flow to define a plurality of corners. A fluid vortex is generated as the fluid flow rolls over each of the corners. A spray ring is disposed along a trailing edge of one of the flowpath structures and is integrally attached to the diffuser. The spray ring includes a plurality of fuel delivery apertures adapted to spray fuel into respective ones of the fluid vortices to form the combustion mixture.

Apparatus for forming a combustion mixture in a gas turbine engine

A cast metallic airfoil structure having a high stiffness core positioned therein for increasing the stiffness of the airfoil. In one embodiment, the high stiffness core is preferably formed of gamma titanium aluminde or a titanium metal matrix composite and is metallurgically bonded to the cast metallic airfoil structure. In one form the airfoil structure forms a portion of an integrally bladed rotor, and in a second form the airfoil structure has a dove tail attachment portion.

High stiffness airoil and method of manufacture

The present invention contemplates a hybrid magnetic bearing thrust disk. In one embodiment the hydrid magnetic bearing thrust disk comprises a high magnetic permeable portion that is supported by a high specific strength circumferential ring. The high specific strength circumferential ring resisting the dynamic loads applied to the hydrid thrust disk.

High speed, high temperature hybrid magnetic thrust bearing

A gas turbine engine has a turbine shroud and back seal plate to define a gas flow passage across turbine rotor blades. A conically shaped turbine support plate for the shroud and back seal plate has a central hub fixedly secured to a bearing mount on the inboard face of a scroll of a compressor for supplying air to a turbine combustor. The conical turbine support plate permits a first predetermined thermal excursion of the shroud and back seal plate to minimize gas leakage between the turbine blades and the turbine shroud components. A plurality of restrictor units are fixedly secured on the scroll and extend axially therefrom to locate a cap portion thereon in close spaced relationship to the turbine support plate to limit expansion of the turbine shroud components with respect to the rotor blades to prevent interference therebetween under elevated temperature conditions.

Growing community of inventors

Greenwood, IN, United States of America

Brian Paul King