Honeywell International Inc.

Derek Anthony Rice

David R Waldman

Devinder N Katariya

Raymond Mccage

Venkatesh S Krishnan

Gary G Capek

Bruce Dan Bouldin

Yates Wong

Michael T Barton

Brian Hann

Kevin Sing

Thomas E Strangman

Christopher Zollars

Richard E Jackson

Amer Aizaz

Andrew F Hieber

Brian Jardine

William C Baker

Paul E Hruska

Daniel Greving

Joseph Becar

Charles P Silcox

Redge How

Maher A Serag

Paul D Rippey

Henry L Chester

Ramasami (Rajan) Thiyagarajan

Hsin-Yi Liu

David A Innes

Gordon B Bailey

Adrienne Lamm

James S Perron

A diffuser and deswirl flow structure includes a plurality of tube structures with an outer wall that is hollow and elongate and that extends between a first portion and a second portion. The plurality of tube structures is disposed in an annular arrangement about the longitudinal axis. The flow structure also includes a plurality of flow passages extending through the tube structures. The plurality of flow passages extend from the first portion to the second portion, respectively. The plurality of flow passages respectfully include a diffuser portion, which is proximate the first portion and configured to diffuse a fluid flow from a compressor wheel. The plurality of flow passages respectfully include a deswirl portion, which is proximate the second portion and configured to deswirl the fluid flow from the diffuser portion. The outer wall defines the diffuser portion and the deswirl portion. The outer wall is self-supporting.

Monolithic diffuser and deswirl flow structure for gas turbine engine

A system for delivering a flow stream of a gas to a compressor. A shroud extends from an inlet to the compressor and defines a main inlet passage configured to direct the flow stream from the inlet to the compressor. A communication plenum is separated from the main inlet passage. A port system includes first and second port subsystems that each provide an opening between the main inlet passage and the communication plenum. The first port subsystem is disposed further from the compressor than the second port subsystem. The port system is configured so that a portion of the gas enters or exits the compressor through the second port subsystem, depending on operating conditions of the compressor.

Ported shroud system for turboprop inlets

A turboprop engine system for an aircraft includes an engine, a propeller, and a gear train coupled to and configured to provide power from the engine to the propeller at a predetermined gear reduction. The engine system also includes a gearbox that houses at least part of the gear train. The gearbox includes a gearbox flow structure and an inlet flow structure that is removably attached to the gearbox. The inlet flow structure and the gearbox flow structure cooperate to define an inlet flow passage to the engine. The inlet flow passage has an upstream end and a downstream end that are cooperatively defined by the inlet flow structure and the gearbox flow structure. The upstream end is configured to receive an airstream that is directed along the inlet flow passage to the downstream end and toward the engine.

Inlet flow structure for turboprop engine

High temperature rolling element bearings and methods for manufacturing high temperature bearing components, such as bearing races or rings, are provided. In one embodiment, the method includes obtaining a powder mixture containing a superalloy powder admixed with hard wear particles, such as carbide particles. The powder mixture is consolidated utilizing a spark plasma sintering process during which the powder mixture is compressed into a sintered blank, while an electrical current is conducted through the powder mixture to heat the powder mixture to a sintering temperature. The sintered blank is then machined to impart the bearing component with its final shape. Precipitate hardening may also be performed, if desired. The spark plasma sintering process is controlled to limit the temperature and duration of the powder consolidation process thereby imparting the resulting bearing component with an enhanced hot hardness and other desirable properties at highly elevated operating temperatures.

Methods for manufacturing high temperature bearing components and rolling element bearings

Methods of forming dual microstructure components include consolidating a powder material comprising an alloy to form a billet, the billet having a first grain structure, inductively heating the billet at an inductive heat treat temperature above a gamma prime solvus temperature of the alloy and subjecting the billet to a subsolvus heat treat temperature that is below the gamma prime solvus temperature of the alloy, waiting a period of time for the first grain structure in an outer portion of the billet to transform into a second grain structure that is coarser than the first grain structure, after the steps of inductively heating and subjecting the billet to the subsolvus heat treat temperature. The methods also include dividing the billet into at least two sections, and machining a final shape into one or more of the at least two sections to form the dual microstructure component.

Methods of forming dual microstructure components

A rotor component assembly including at least one void space formed therein that provides for reduction in weight of the component and/or cooling of the component during operation. The rotor component is formed by positioning a coated mild steel insert within a mold having an internal cross-section substantially the same in dimensions as the final rotor component. The mold is filled with a superalloy metal powder and undergoes hot-isostatic pressing to consolidate the powder about the coated core insert and form a superalloy structure. The mold is removed from about the superalloy structure and the core insert is removed from within the superalloy structure, thereby defining the at least one internal void within the rotor component.

Superalloy rotor component and method of fabrication

There is provided a method for fabricating a dual alloy structure that may in turn be machined to fabricate a rotary component for use in a gas turbine engine. The method provides a powder metal (PM) nickel based superalloy and a nickel aluminide intermetallic based alloy. The powder metal (PM) nickel based superalloy displays characteristics, such as improved strength, low cycle fatigue resistance, fracture toughness, and crack growth resistance. The nickel aluminide intermetallic based alloy displays characteristics, such as high temperature creep and oxidation resistance, suitable for use in the outer radial area of an impeller. A bore sub-element is fabricated from the powder metal (PM) nickel based superalloy. A body sub-element is fabricated from the nickel aluminide intermetallic based alloy. The bore sub-element and body sub-element are joined by inertia welding or diffusion bonding at a common mating surface.

Method of manufacture of a dual alloy impeller

A dual alloy turbine rotor is provided and includes an integrally formed blade ring and a disk. The blade ring is made of a first alloy and includes a ring portion and a plurality of airfoils extending therefrom, where the ring portion includes an inner surface, and each airfoil includes an internal cavity formed therein. The disk is made of a second alloy and may have a beveled outer peripheral surface. The disk is disposed within the ring portion such that the disk outer peripheral surface contacts at least a portion of the ring portion inner surface. Methods of manufacturing the turbine rotor are also provided.

Multi-alloy turbine rotors and methods of manufacturing the rotors

A coating system is used on an engine component having an outer surface configured to be exposed to a first plurality of particles impinging against the outer surface at an angle within a first angle range and a second plurality of particles impinging at an angle in a second angle range. The system includes a bond layer overlying the engine component outer surface, a first erosion-resistant layer comprising a first material that is more resistant to erosion by particles impinging the component outer surface at an angle within the first angle range than by particles impinging within the second angle range, an interlayer overlying the first erosion-resistant layer, and a second erosion-resistant layer comprising a second material that is more resistant to erosion by particles impinging the component outer surface at an angle within the second angle range than by particles impinging within the first angle range.

Multilayered erosion resistant coating for gas turbines

A spun metal form used to manufacture a dual alloy turbine wheel. The spun metal form prevents braze alloy contamination of the hub/casting interface by relocating the braze joint away from the interface. The spun metal form allows for the re-working of components after failed vacuum brazing and increases the time to failure of dual alloy turbine wheels.

Spun metal form used to manufacture dual alloy turbine wheel

A hoop stress relief mechanism is disclosed for use on a rotary body to relieve stress caused by both thermal and centrifugal forces. The mechanism may consist of a J-shaped slot cut from the outer rim of the rotary body a distance inwardly toward the axis of rotation, the slot having a curve in its inward end that curves back towards the outer rim. The J-shaped slot may extend through the rotary body to join its two faces. The J-shaped slot may be fabricated by an electric discharge wire machine. The electric discharge wire machine may make multiple passes in order to smooth the bottom surface of the curved slot portion.

Hoop stress relief mechanism for gas turbine engines

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Phoenix, AZ, United States of America

Derek Anthony Rice