Abstract: A combustor for a gas turbine engine including a combustor liner support shell with a furrow formed therein; a forward liner panel mounted to the support shell via a multiple of studs, the forward liner panel including a forward liner panel rail the extends into the furrow; and an aft liner panel mounted to the support shell via a multiple of studs downstream of the forward liner panel, the aft liner panel including an aft liner panel rail that extends into the furrow.

Abstract: A liner of a combustor wall assembly generally defines a combustion chamber and includes a film cooling circuit having a channel communicating through a hot face of a substrate of the liner. An aperture of the circuit is defined by the substrate, extends through the cold face and is in fluid communication with the channel. The hot face of the substrate is covered with a coating that extends over and thus defines in-part the channel. A film cooling hole extends through the coating and is in fluid communication with the channel. A method of manufacturing the circuit includes casting the substrate with the aperture and hole; then placing an insert into the channel prior to application of the coating over the substrate and insert. The insert is then removed and the film cooling hole is formed through the coating.

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: An additive manufacturing method includes segmenting a CAD file of a component along a 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: A combustor for a gas turbine engine including a forward liner panel mounted to a support shell via a multiple of studs, the forward liner panel including an aft non-linear circumferential edge and an aft liner panel mounted to the support shell via a multiple of studs, the aft liner panel including a forward non-linear circumferential edge that is complementary to the aft non-linear circumferential edge.

Abstract: A combustor of a gas turbine engine includes a combustor shell having a diffuser side facing a diffuser chamber and a combustor side facing a combustor chamber. The combustor may include a first combustor panel coupled to the combustor side of the combustor shell and a second combustor panel coupled to the combustor side of the combustor shell. A gap may be defined between the first combustor panel and the second combustor panel and the combustor shell may include a gap impingement hole that is directly open to and is configured to deliver cooling air directly to the gap. In various embodiments, the combustor further includes a throttle plate coupled to the diffuser side of the combustor shell.

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: 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: An apparatus includes a component built by layer-by-layer additive manufacturing. An external structure is located on an external surface of the component. An internal structure is positioned within the component and is integrally formed to the external structure. The internal structure is made of a matrix structure, honeycomb structure, or lattice structure. The internal structure provides structural support, vibration dampening, heat transfer, energy absorption, fluid flow, or piping for the component.

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: 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: Gas turbine engine systems and methods involving enhanced fuel dispersion are provided. In this regard, a representative method for operating a gas turbine engine includes: providing a gas path through the engine; introducing a spray of fuel along the gas path downstream of a turbine of the engine; and impinging the spray of fuel with a relatively higher velocity flow of air such that atomization of the fuel is increased.

Abstract: A combustor for a gas turbine engine including a liner panel mounted to a support shell via a multiple of studs, the liner panel including a forward section and an aft section that defines an arcuate surface section therebetween in the axial profile between the forward section and the aft section.

Abstract: A combustor for a gas turbine engine including a support shell with a shell coating applied thereto and a multiple of liner panels circumferentially mounted within the support shell via a multiple of studs, each of the multiple of liner panels having a liner coating applied thereto.

Abstract: A combustor for a gas turbine engine including a forward liner panel mounted to a support shell via a multiple of studs, the forward liner panel including an aft non-linear circumferential edge and an aft liner panel mounted to the support shell via a multiple of studs, the aft liner panel including a forward non-linear circumferential edge that is complementary to the aft non-linear circumferential edge.

Abstract: A combustor for a gas turbine engine including a liner panel mounted to a support shell via a multiple of studs, the liner panel including a forward section and an aft section that defines an angle therebetween in the axial profile between the forward section and the aft section.

Abstract: A combustor for a gas turbine engine including a combustor liner support shell with a furrow formed therein; a forward liner panel mounted to the support shell via a multiple of studs, the forward liner panel including a forward liner panel rail the extends into the furrow; and an aft liner panel mounted to the support shell via a multiple of studs downstream of the forward liner panel, the aft liner panel including an aft liner panel rail that extends into the furrow.

Abstract: An ignition system for a combustor of a gas turbine engine is disclosed. The ignition system may include an igniter operatively associated with the combustor, and an electrode operatively associated with the combustor and spaced from the igniter, wherein an electrical potential is created between the igniter and the electrode to produce an electric arc therebetween.

Abstract: A combustor of a gas turbine engine includes a combustor shell having a diffuser side facing a diffuser chamber and a combustor side facing a combustor chamber. The combustor may include a first combustor panel coupled to the combustor side of the combustor shell and a second combustor panel coupled to the combustor side of the combustor shell. A gap may be defined between the first combustor panel and the second combustor panel and the combustor shell may include a gap impingement hole that is directly open to and is configured to deliver cooling air directly to the gap. In various embodiments, the combustor further includes a throttle plate coupled to the diffuser side of the combustor shell.

Abstract: A combustor for a gas turbine engine including a non-linear axial edge between the forward edge and the aft edge. A combustor for a gas turbine engine including a multiple of forward liner panels circumferentially mounted within the support shell via a multiple of studs, each of the multiple of forward liner panels having a non-linear axial edge therebetween to define a forward non-linear interface between each adjacent pair of the multiple of forward liner panels and a multiple of aft liner panels circumferentially mounted within the support shell via a multiple of studs aft of the multiple of forward liner panels, each of the multiple of aft liner panels having an aft non-linear axial edge therebetween to define a non-linear interface between each adjacent pair of the multiple of aft liner panels.