United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Mitsubishi Heavy Industries America, Inc.

Hoyt Y Chang

Shelton O Duelm

Shayan Ahmadian

David A Burdette

Michael G McCaffrey

Thomas N Slavens

Stephen Karl Kramer

Joshua Daniel Winn

Stanislav Kostka, Jr

Brooks E Snyder

Timothy Dale

Christos Adamopoulos

Matthew B DeGostin

Yusuke Ichihashi

Yohei Fujimoto

Brooke E Snyder

Sheryan Ahmadian

Jr Stanislav Kostka

A method for implementing a component design model includes defining a component design specification, defining at least one variation mode parameter of a component manufacturing process, constructing a parametric model relating the at least one variation mode parameter to the component design specification, determining an expected component output based on the parametric model and comparing the expected component output to at least one as-manufactured component, and defining the parametric model as accurate in response to the expected output component matching the at least one as-manufactured component within a predefined degree of accuracy.

Parametric component design process

A self-contained computerized system includes a processor and a memory. The memory stores instructions for causing the self-contained computerized system to perform the process of morphing a finite element mesh of a legacy design to the finite element mesh of a component design and applying measurement variations from the legacy design to the finite element mesh of the component design.

Process for gas turbine engine component design

A coating system including a reflective cool down chamber with at least one arcuate wall; and an infrared lamp directed at the arcuate wall.

System and method for low thermal shock-fast cooling of thermal barrier coating

One exemplary embodiment of this disclosure relates to a gas turbine engine including a plate, a frame attached to the plate, and a panel. The panel is supported by the frame.

Protective panel and frame therefor

A heat shield for use in a combustor of a gas turbine engine including a plurality of standoff pins that extend from a cold side, the plurality of standoff pins at least partially surround an attachment stud, the cold side including at least one film cooling hole adjacent to the plurality of standoff pins and the attachment stud.

Additive manufactured combustor heat shield

A case is provided for a gas turbine engine. The case includes a wall defining a through-hole. The case also includes first and second pockets adjacent to, and on opposite sides, of the through hole. A method of reducing stress in a case of a gas turbine engine is also provided that includes reducing stress about a through-hole by providing a concavity on each side of the through-hole.

Shielding pockets for case holes

A gas turbine engine ceramic panel assembly includes a ceramic liner that has a slot and includes a hole. An insert is received in the slot and provides a female fastening element aligned with the hole. A method of manufacturing a ceramic panel assembly includes the steps of forming a ceramic liner with a slot, installing an insert into the slot, and securing the ceramic liner to a structure using a fastening assembly.

Gas turbine engine with ceramic panel

A combustor includes a liner and a support. The liner includes first and second spaced apart liner walls defining a combustion chamber therebetween, and an end wall joining the first and second liner walls at a forward end. The support is configured to resiliently support the liner and includes first and second spaced apart support walls having first and second retainers configured to capture aft ends of the first and second liner walls. A gas turbine engine and a method of assembling a combustor are also disclosed.

Combustor liner

A vane assembly within a mid-turbine frame of a gas turbine engine includes an airfoil that extends between an outer platform and an inner platform. The airfoil includes an outer wall defining a leading edge of a first radius. An inner wall of the airfoil defines an inner cavity including a forward portion proximate the leading edge defining a second radius different than the first radius.

Stress mitigation feature for composite airfoil leading edge

A cooled wall for lining a combustor of a gas turbine engine includes an annular outer wall defining a radial passage extending therethrough, an inner wall spaced radially from the outer wall to define an axial channel in fluid communication with the radial passage, and a connecting wall joining the inner wall to the outer wall. A method of making the cooled wall includes providing a three-dimensional computer model of the cooled wall, depositing a uniform thickness layer of material on a substrate, using a laser or electron beam to melt or sinter the material to form a cross section, repeating the depositing and cross section steps to form the cooled wall, and heat treating the cooled wall.

Combustor wall cooling channel formed by additive manufacturing

A combustor thermal shield is provided. The combustor thermal shield may have a combustor panel that may protect a surface of a combustor from heat and/or flame. The combustor thermal shield may be mounted to the combustor by an attachment stud formed as a unitary body with the combustor panel. The combustor panel and the attachment stud may be made of different materials. A transition region may be disposed between the attachment stud and the combustor thermal shield and formed as a unitary body with the combustor panel and the attachment stud. The transition region may be made of a mixture of the different materials. The mixture may vary according to a gradient as a function of proximity to the combustor panel and the attachment stud.

Combustor panel having material transition region

A thermal mechanical dimple array for combustor liners of gas turbine engines are located at known hot spots along liner to reduce stress and minimize cracking. Such arrays have indentations in the liners at the hot spots that act to thermal mechanically flex the liner at the hot spot away from the heat source and distribute stresses such that cracking of the liner is reduced and durability is improved.

Thermal mechanical dimple array for a combustor wall assembly

A combustor is provided. The combustor may comprise a shell including a ceramic material. A gradient transition from ceramic to metal may have a predominantly ceramic side coupled to the shell. A metal mating surface may also be coupled to a predominantly metal side of the gradient transition from ceramic to metal.

Ceramic and metal engine components with gradient transition from metal to ceramic

A fairing assembly for a gas turbine engine includes an inner fairing platform; an outer fairing platform located radially outward of the inner fairing platform; a plurality of vanes radially outward from the inner fairing platform between the inner fairing platform and outer fairing platform; a mount located on a radially outer side of the outer fairing platform, the mount attached to the outer fairing platform on the radially outer side; and a link that includes an elongated member with a first end and an opposite second end, the first end configured to fit within the mount and define a pivot axis of the link with respect to the mount.

Connection for a fairing in a mid-turbine frame of a gas turbine engine

An airfoil for a gas turbine engine has a first layer forming a cavity having transitioning from a first thickness to a second thickness through a ply drop region. A second layer is secured to the first layer.

Growing community of inventors

Manchester, CT, United States of America

Hoyt Y Chang