Raytheon Technologies Corporation

Alstom Technology Limited

Chromalloy Gas Turbine Corporation

United Technologies Corporation

Ansaldo Energia IP UK Limited

Adam Lee Hart

William S Pratt

Luis Miguel Ayalde

Brian Levy

Jose R Paulino

Jonathan Jeffery Eastwood

Brian Richard Craig

Daniel L Folkers

W David Day

Ryan M Stanley

Matthew E Bintz

Vincent C Martling

Zhenhua Xiao

Steven D Roberts

Charles A Ellis

James Page Strohl

Douglas James Dietrich

Stephen Wayne Fiebiger

Orestes J Maurell

Richard Ramirez

John Cuyler Hunt

Gerald Raresheid

Elliot I Goodman

Kuo-Ting Hsia

Samer Abdel-Wahab

Adam John Fredmonski

Nicolas Bueno

Gaurav K Joshi

Esref Yaylagan

Patrick Scheelk

A blade outer air seal (BOAS) for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion; and a multitude of cooling fins disposed around the radially outer face of the seal body, at least one of the multitude of cooling fins having at least one aperture.

Blade outer air seal cooling fin

An apparatus and method for mounting a combustion liner within a flow sleeve of a gas turbine combustion system is disclosed. A mounting system comprises a plurality of low-profile mounting tabs secured to a combustion liner where each of the mounting tabs are placed within slots of flow sleeve pegs when the combustion liner is installed in a flow sleeve. A plurality of liner stop brackets are removably secured to a flange of the flow sleeve and have an arm extending to be adjacent to a top contact surface of the mounting tabs. The mounting system reduces blockage to the surrounding airflow.

Axial stop configuration for a combustion liner

A synchronizing ring assembly includes a synchronizing ring portion that has a first and a second distal end, the first and the second distal end each form an integrated surge bumper.

Synchronizing ring surge bumper

A vane arm assembly for a gas turbine engine is provided including: a vane arm having a first end, a second end opposite the first end, and an aperture proximate the second end, the aperture being defined by an aperture wall; a vane stem extending through the aperture of the vane arm; a mechanical fastener retaining a position of the vane arm in the longitudinal direction of the vane stem; and an impedance clip partially enclosing a portion of the second end of the vane arm to provide redundant position retention of the vane arm in the longitudinal direction of the vane stem.

Continuous wedge vane arm with failsafe retention clip

A system for retaining a cross fire tube in a multi-combustor gas turbine engine is disclosed. The system comprises a flow sleeve having a generally annular body, and a flange at a forward end thereof and having one or more recessed portions. A cross fire tube extends through one or more openings in the flow sleeve and is secured in place by a retention clip. The retention clip includes a plurality of fingers which engage the cross fire tube and a mounting plate engaging the one or more recessed portions of the flow sleeve flange so as to create a clip engagement having a lower profile than prior art configurations.

Cross fire tube retention system

A synchronizing assembly including: a synchronizing ring having a countersunk orifice; and a bumper assembly comprising: a bumper at least partially enclosing a cavity, the bumper comprising an outward side, an inward side opposite the outward side, a first side extending between the outward side and the inward side, a first opening on the outward side extending into the bumper to define a first portion of the cavity; and a second opening on the first side extending into the bumper to define a second portion of the cavity; and a bolt comprising a bolt head and a bolt shank, wherein the bolt shank extends through the first opening and the bolt head is secured within the second portion of the cavity, wherein the bumper further comprises a raised boss extending away from the outward surface, the raised boss configured to mate with the countersunk orifice of the synchronizing ring.

Integral shear locking bumper for gas turbine engine

A heat shield for a gas turbine engine includes a radial heat shield flange that extends in a circumferential direction and forms a ring. A plurality of bosses extend from a first axial side of the radial heat shield flange. There is a plurality of guide pins. One of the plurality of guide pins extends from a corresponding one of the plurality of bosses.

Gas turbine engine exhaust component

An example case for a gas turbine engine includes a bearing compartment, a radially outer ring, and a radially inner ring. The radially outer ring and radially inner ring are connected by a plurality of circumferentially spaced apart struts and define an annular flow path therebetween. The radially inner ring includes a first flange that extends radially outward from the bearing compartment. The first flange includes an undulation that extends axially away from the plurality of struts, and defines a load path between the bearing compartment and the radially outer ring.

Gas turbine engine case including bearing compartment

A gas turbine engine component has an axial seal pin assembly for sealing between adjacent platform regions where the seal pin assembly design provides a seal pin damper slot extending to a forward axial stop in a forward buttress region. The forward axial stop includes a relief slot to interrupt the loadpath from the airfoil into the blade neck and attachment regions to reduce the stress concentrations in these regions.

Axial retention of a platform seal

A turbine blade, airfoil configuration, and rotor stage are disclosed in which through the airfoil profile disclosed in Table 1, a modification in airfoil flutter and swirl are achieved. Through the airfoil configuration, the swirl and improved platform sealing configuration, result in the turbine blades having the airfoil profile with an increased performance output from the turbine.

Turbine blade having improved flutter capability and increased turbine stage output

A compressor diaphragm for a gas turbine engine having improved wear capability, manufacturability, and assembly techniques is disclosed. The diaphragm includes a shiplap-type joint at an outer vane platform for connecting to adjacent vane assemblies and a clamshell-like assembly of a seal box secures and seals regions around the inner vane platform of the compressor diaphragm so as to reduce wear between the seal box and the vane assemblies. The inner platform of the diaphragm segments are fastened to each other through circumferentially-oriented fasteners at the inner diameter platform.

Compressor vane diaphragm

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Stuart, FL, United States of America

Adam Lee Hart