United Technologies Corporation

Raytheon Technologies Corporation

Jonathan Perry Sandoval

Scott Daniel Virkler

Julian Partyka

Edwin Otero

Daniel A Snyder

Kyle C Lana

Randolph J Smith

Carey J Clum

Russell J Bergman

Joseph T Caprario

Jonathan P Burt

Raymond S Hummel

Jaimie Schweitzer

Jaimie Taraskevich

Pamela A Humphrey

Kyle Charles Lana

A combustor wall assembly has a heat shield and a supporting shell with a cooling cavity defined therebetween. A grommet generally includes a wall defining a dilution hole isolated from the cooling cavity, and a flange projecting radially outward from the wall and into the cooling cavity. The flange is space from the heat shield and a cooling channel is defined between the wall and the heat shield that communicates with the cavity for cooling the wall proximate to a combustion chamber.

Cooled grommet for a combustor wall assembly

An assembly for a gas turbine engine includes a minidisk that includes an axial extension extending from a disc. The axial extension includes an inner diameter surface and a recess arranged radially opposite the inner diameter surface. The recess provides a radially outwardly extending flange and a bumper extending radially inward from and proud of the inner diameter surface. A method of working on a gas turbine engine section includes inserting a tool into a cavity beneath a seal assembly, and engaging a flange of a minidisk with the tool to manipulate first and second rotors with respect to one another.

Turbine minidisk bumper for gas turbine engine

An example turbine includes a turbine disk and a turbine blade. The turbine disk includes a plurality of lugs and a plurality of slots. Each of the plurality of lugs is located between two of the plurality of slots. Each of the plurality of lugs includes a tab that extends radially outwardly from an end of the lug. The tab has a first section having an upper surface and a second section having an upper surface. The upper surface of the first section is inclined greater than the upper surface of the second section, and the upper surface of the first section of the tab defines a contour. The turbine blade includes a root received in one of the plurality of slots and a platform, and the platform has a lower surface defining a contour. When the turbine blade is received in one of the plurality of slots, the lower surface of the platform is located above a portion of the upper surface of the first section of the tab, and the contour of the upper surface of the first section of the tab matches the contour of the lower surface of the platform.

Turbine blade damper seal

A curvic seal for a gas turbine engine includes a cylindrical seal with a retention feature.

Curvic seal for gas turbine engine

According to an exemplary embodiment of this disclosure, among other possible things a gas turbine engine includes a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor section, a plurality of turbine disks in at least one of the compressor section and the turbine sections, at least one cover plate corresponding to at least one of the turbine disks, each of the cover plates includes, at least two snaps connected via a webbing portion, and a bore region radially inward of the at least two snaps and connected to at least one of the at least two snaps via the webbing portion.

Double snapped cover plate for rotor disk

A combustor component defining a dilution hole configuration for a combustor wall assembly of a gas turbine engine has first and second dilution holes that communicate radially with one-another and with respect to respective centerlines of each dilution hole. In operation, each dilution hole forms a respective jet stream along respective centerlines for flowing cooling air into a combustion chamber.

Multi-streamed dilution hole configuration for a gas turbine engine

A rotor assembly includes a first rotor, a second rotor mounted on the first rotor and co-rotatable there with, and a thermal shield interlocked with the second rotor for co-rotation there with.

Interlocking rotor assembly with thermal shield

A thermally compliant grommet assembly for a combustor includes first and second shells spaced apart from one another and a grommet having a tubular portion projecting into a hole in the second shell and defining a bore in fluid communication through the first and second shells. The grommet further has an oblong flange portion extending circumferentially around and projecting outward from the tubular portion. The oblong flange portion is located between and is in sealing contact with the first and second shells. At least one of the first and second shells is constructed and arranged to move between cold and hot states and generally along a thermal displacement path. The oblong flange portion is generally elongated along the path to maintain sealing contact with at least one of the shells as the shells move between the hot and cold states.

Thermally compliant grommet assembly

A damper seal received in a cavity of a turbine blade located between a platform and a retention shelf damper seal according to an exemplary aspect of the present disclosure includes, among other things, a central body portion having a first end region, an opposing second end region, and a width. The damper seal further includes a first portion extending from the first end region of the central body portion, and a first end region of the first portion includes first outwardly extending tabs that define a first enlarged portion that has a first width greater than the width of the central body portion and a second portion extending from the opposing second end region of the central body portion.

A disk-seal arrangement for a gas turbine engine rotor includes a blade retention disk having a longitudinally opening slot therein and a seal disk juxtaposed to said blade retention disk and a single anti-rotation tab received within the blade retention disk slot. The seal disk also includes a pair of balance slots disposed immediately adjacent the antirotation tab to offset the weight thereof for preserving rotor balance. A split ring is disposed longitudinally between the blade retention and seal disks for reacting longitudinal loading therebetween. The ends of the split ring seat against side surfaces of the antirotation between the split ring and the blade retention and seal disks.

Gas turbine engine rotor disk-seal arrangement

In accordance with one aspect of the disclosure, a rotor is disclosed. The rotor may include a disk having a central axis, an airfoil radially extending from the disk, a bayonet tab extending radially from the disk, and a lock. The lock may further include a short tab and a long tab, both extending radially from the disk and in an axial direction with respect to the central axis. The long tab may have a greater axial length than the short tab.

Lock for retaining minidisks with rotors of a gas turbine engine

A conjoined grommet assembly for a combustor wall assembly of a gas turbine engine has a first grommet defining at least in-part a first dilution hole, and a second grommet defining at least in-part a second dilution hole. The first and second dilution holes are spaced closely together such that the first grommet is in contact with the second grommet.

Conjoined grommet assembly for a combustor

A gas turbine engine combustor configured to cool a first vane platform by directing airflow through an aft combustor rail is disclosed. In various embodiments, a gas turbine engine combustor may comprise an aft combustor rail connected between an outer wall of an outer liner and an aft heat shield panel, and a plurality of holes in the aft combustor rail. The aft combustor rail may be positioned near an outer vane platform of a first vane stage of a turbine. Further, the plurality of holes may be designed to focus pass-through airflow onto the outer vane platform. In various embodiments, each hole of the plurality of holes may be angled between 20&#xb0;-90&#xb0; relative to the aft combustor rail. Additionally, the plurality of holes may be angled to converge the airflow to a focused area, such as a leading edge vane airfoil of the first vane stage.

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Hebron, CT, United States of America

Jonathan Perry Sandoval