Abstract: A seal for use between a first duct and a second duct and, the ducts having relative motion therebetween, includes a first plurality of fingers for attachment to the first duct to be disposed about one of the first duct and the second duct, a seal land attaching to about an other of the first duct and the second duct, and a seal attaching to the seal land and in contact with the plurality of fingers wherein one of the plurality of fingers and the seal is metallic and an other of the seal land or the plurality of fingers is non-metallic.

Abstract: A unitary one-piece hub has first and second rings and a midsection arranged between the first and second rings. The midsection includes a plurality of windows configured to receive a plurality of cross members. The windows include a lip configured to surround the cross members. A gas turbine engine and a method of providing a hub for a gas turbine engine are also disclosed.

Abstract: A gas turbine engine has a first component and a second component. The first and second components have a high-pressure chamber on one side and a low pressure chamber on an opposed side. A three sided seal has one side facing each of the first and second components, and a third side facing a third component. At least one non-metallic wear surface is between one of the three sides of the seal and the facing component.

Abstract: An assembly may comprise a first part including a first surface, and a second part including a second surface and a bottom edge having sloped regions and at least one well between the sloped regions. The assembly may further comprise a brazed joint joining the first surface to the second surface. The at least one well may be configured to collect excess braze material from the brazed joint.

Abstract: The present disclosure relates generally to a system for fan nacelle inlet flow control in a gas turbine engine, the nacelle comprising a nacelle inlet cowl including an inlet lip disposed at a leading edge of the nacelle inlet cowl, an inner surface extending aft from the inlet lip, and an outer surface extending aft from the inlet lip and positioned radially outward of the inner surface; and at least one flow control passage extending through the nacelle inlet cowl, each of the at least one flow control passage including a flow control passage inlet, disposed on the inlet lip, and a flow control passage outlet; wherein air may flow into the flow control passage inlet, through the flow control passage, and exits the flow control passage outlet.

Abstract: A component for a gas turbine engine includes a substrate with a substrate aperture and a coating on the substrate that extends a length of the substrate aperture. A liner assembly for a gas turbine engine includes a hot sheet with a multiple of apertures and a coating on the hot sheet that extends a length of each of the multiple of apertures. A method of forming an aperture to provide film cooling in a component of a gas turbine engine, includes forming a multiple of substrate apertures in a substrate. Each of the multiple of substrate apertures defines a substrate inner periphery. A coating is applied on the substrate after forming the multiple of substrate apertures to define a coating inner periphery at least partially within each of the multiple of substrate apertures. The coating inner periphery is smaller than the substrate inner periphery.

Abstract: A gas turbine engine component and a method for forming a component with a plurality of apertures are provided. An additive metal manufacturing process for fabricating a component with apertures includes receiving data including a three-dimensional representation of the component, generating an electronic file based on the received data, wherein the electronic file includes fabrication instructions for entire portions of the component, and forming initial and additional portions of the component based on the electronic file.

Abstract: A component is provided for a gas turbine engine includes a substrate with an aperture. The component also includes a backside coating on a backside of the substrate and at least partially onto an inner boundary of the aperture, where the backside coating forms a passage with the aperture. A method of forming a shaped aperture in a component of a gas turbine engine is provided. The method includes applying a backside coating on a backside of a substrate and at least partially onto an inner boundary of an aperture. The backside coating forms a passage with the aperture including a convergent section, a divergent section and a throat therebetween.

Abstract: A method includes building an article by a layer-by-layer additive manufacturing process. While the article is being built, a solid outer wall is formed. An inner structure of the article is integrally formed with the outer wall. The inner structure includes an internal permeable structure.

Abstract: An assembly may comprise a first part including a first surface, and a second part including a second surface and a bottom edge having sloped regions and at least one well between the sloped regions. The assembly may further comprise a brazed joint joining the first surface to the second surface. The at least one well may be configured to collect excess braze material from the brazed joint.

Abstract: In accordance with one aspect of the disclosure, a swirler is disclosed. The swirler may include an outer shroud and inner shroud. The inner shroud may be positioned radially inside the outer shroud. At least one of the outer shroud and inner shroud may have a major diameter which is larger than a minor diameter such that the shrouds define an oblong shape. The swirler may further include a plurality of vanes which may be positioned between the inner and outer shrouds.

Abstract: A combustor for a gas turbine engine includes an forward fuel injection system in communication with a combustion chamber and a downstream fuel injection system that communicates with the combustion chamber downstream of the forward fuel injection system.

Abstract: Aspects of the disclosure are directed to a liner associated with a combustor of an aircraft engine, comprising: a thermal barrier coating, and a base metal, wherein the thermal barrier coating comprises a contoured surface on a flowpath side proximate to an exit of a hole formed by the thermal barrier coating and the base metal.

Abstract: A combustor includes a combustion chamber and a combustor wall that at least partially bounds the combustion chamber. The combustor wall includes a first side that faces the combustion chamber and a second side that faces away from the combustion chamber. The second side includes an array of geometric surface features for thermal transfer. A metallic coating is disposed on at least a portion of the second side of the combustor wall.

Abstract: A swirler includes an inner shroud positioned radially inside an outer shroud. At least one of the outer shroud and inner shroud has a major diameter and a minor diameter, the major diameter being greater than the minor diameter, the major and minor diameters defining an ovate shape. The swirler further includes a plurality of vanes extending between the inner and outer shrouds.

Abstract: Systems and methods are described herein whereby an air jet is configured to manipulate local aerodynamics and/or boundary layer flows associated with a dilution hole. A gas turbine component including a combustor panel, a dilution hole located within the combustor panel and an air jet located within the combustor panel positioned in close proximity to the dilution hole is described. The dilution hole is configured to produce a flow of cooling fluid. An air flow from the air jet is configured to deflect secondary flows produced within a combustor. The air jet is located close enough to a leading edge of the dilution hole such that the air flow from the air jet manipulates a pressure gradient of the dilution hole.

Abstract: A wall assembly that may be for a combustor of a gas turbine engine includes a liner having a hot face that defines a combustion chamber, an opposite cold face, and a plurality of effusion holes. A shell of the assembly is spaced outward from the cold face and includes a plurality of impingement holes each having a centerline orientated substantially normal to the cold face. A plurality of cooling member arrays of the liner each include a first plurality of members that may be pins projecting outward from the cold face to conduct heat out of the liner. Each array is spaced between adjacent effusion holes and is symmetrically orientated about the respective centerline.

Abstract: An additive manufacturing method includes segmenting a CAD file of an component along an build interface to define at least a first component segment and a second component segment each of the first component segment and the second component segment sized to fit within an additive manufacturing build chamber; additive manufacturing the first component segment and the second component segment within the build chamber; and bonding the first component segment and the second component segment to form the component.

Abstract: An additive manufacturing system includes a support structure defining a cavity. A movable platform is contained within the cavity, and is capable of moving along a build direction within the cavity. The movable platform is shaped to receive and hold a substrate that includes at least a portion of a finished part. The movable platform, in combination with the substrate, defines a working surface for building a finished product via additive manufacturing.

Abstract: A gas turbine engine has a first component and a second component. The first and second components have a high-pressure chamber on one side and a low pressure chamber on an opposed side. A three sided seal has one side facing each of the first and second components, and a third side facing a third component. At least one non-metallic wear surface is between one of the three sides of the seal and the facing component.