United Technologies Corporation

Other

Raytheon Technologies Corporation

Rtx Corporation

Kevin N McCusker

Matthew E Bintz

William S Pratt

Charles C Wu

Weston Behling

Venkataramana Vijay

Erica J Harvie

Charles A Ellis

David M Nissley

Roger E Paolillo

John H Hughes

Moon-Kyoo Brian Kang

Page Russell Palmiter

Lon M Stevens

Mark S Turner

A rotor stack for a gas turbine engine includes a first rotor disk with a first rotor spacer arm, the first rotor spacer arm having a first flange with an outboard flange surface and an inboard flange surface, a first hole along an axis through the first flange, the first hole having a counterbore in the outboard flange surface; a second rotor disk with a web having a second hole along the axis; a third rotor disk with a third rotor spacer arm, the third rotor spacer arm having a third flange with an outboard flange surface and an inboard flange surface, a third hole along the axis through the third flange, the third hole having a counterbore in the inboard flange surface; and a bushing with a tubular body and a flange that extends therefrom, the tubular body comprising at least one axial groove along an outer diameter thereof, the bushing extends through the first hole, the second hole and partially into the counterbore in the inboard flange surface of the third hole.

Rotor stack bushing with adaptive temperature metering for a gas turbine engine

According to an embodiment disclosed herein, an apparatus for cooling a rotating part having cooling channels therein, the rotating part attaching to a disk rotating about an axis, the disk having a conduit for feeding a cooling fluid to the cooling channel is described. The apparatus has a first impeller rotating with the disk and in register with the conduit and an outer periphery of the disk, the impeller directing the cooling flow to the conduit.

High solidity and low entrance angle impellers on turbine rotor disk

A gas turbine engine is provided with turbine sealing structures including knife edge seals which extend at an angle relative to an axial center line of the engine. Each knife edge seal is associated with a control pocket defined between a radially inner surface and a spaced radially outer surface. The control pockets and their associated knife edge seals create a difficult flow path to prevent leakage into radially inner portions of the turbine section.

Gas turbine engine with canted pocket and canted knife edge seal

A gas turbine engine comprises a compressor section, a combustion section disposed downstream from the compressor section, and a turbine section disposed downstream from the combustion section. The turbine section includes a turbine disk defining a plurality of turbine disk slots for accommodating turbine blades. The plurality of turbine disk slots each include an inlet having a rounded periphery at a bottom portion thereof.

Gas turbine disk slots and gas turbine engine using same

A total condensed circuit for cardiopulmonary bypass and cardioplegia are provided, as well as methods of using the same. The total circuit includes a cardiopulmonary bypass portion including tubing and components which together have a substantially short path length and priming volume preferably under 800 ml. Thus, the opportunity for an inflammatory response caused by blood contacting plasticizers is minimized. The bypass circuit includes a shunt which bypasses a blood reservoir of the total circuit. The cardioplegia circuit infuses cardioplegia fluid into blood pulled from an oxygenator of the bypass circuit. According to the method, either or both of heart-lung bypass and cardioplegia can be performed with only minimal or no isotonic priming solution circulated into the patient. In addition, use of the shunt and reservoir together eliminate the possibility of air entering the circulation system upon kinking of the circuit.

Methods of using condensed perfusion circuit for cardiopulmonary bypass and cardioplegia

Condensed perfusion circuit for cardiopulmonary bypass and cardioplegia

A turbine shroud segment includes a substrate and a coating layer having varying thickness. Various construction details are developed that provide minimal spalling of the coating layer during use of the shroud segment. In a particular embodiment, a shroud segment includes a coating layer that tapers towards the edges. The thickness tapers to a minimum thickness along the leading and trailing edges. Within the blade passing region of the shroud segment, the coating layer tapers towards the lateral edges to a thickness determined by the minimum thickness required for abrasive contact between the shroud segment and rotor blades. In another particular embodiment, the varying thickness of the coating layer is produced by forming the substrate with a concave surface, applying the coating, and subsequently machining back the coating layer to the desired dimensions.

Turbine shroud segment including a coating layer having varying thickness

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West Hartford, CT, United States of America

Kevin N McCusker