A method of preventing spallation for a geometrically segmented thermally insulating top coat on an article, the method comprising forming a surface feature in a surface of the article; forming a crack deflection feature in the surface proximate the surface feature; disposing the thermally insulating topcoat over the surface feature; and forming segmented portions that are separated by faults extending through the thermally insulating topcoat from the surface feature.

A liner assembly includes a core and a septum coupled to the core. The liner assembly also includes a facesheet coupled to the septum. The facesheet includes a plurality of slots defined therethrough. Each slot of the plurality of slots is substantially continuous over a selected portion of a surface and perpendicular to or parallel to a centerline of the liner assembly.

In one exemplary embodiment of the present disclosure a gas turbine engine defining a radial direction and an axial direction is provided. The gas turbine engine includes a stationary frame, a compressor and a turbine, the compressor or the turbine including a first plurality of rotor blades and a second plurality of rotor blades, the first plurality of rotor blades and second plurality of rotor blades alternatingly spaced along the axial direction and rotatable with one another; and a gearbox including a first gear coupled to the first plurality of rotor blades, a second gear coupled to the second plurality of rotor blades, and an intermediate gear positioned between the first gear and the second gear and coupled to the stationary frame, the intermediate gear defining an axis of rotation, the axis of rotation defining an angle with the radial direction less than about 75 degrees.

A gas turbine engine with an accessory gearbox. The accessory gearbox is rotatably coupled to the engine drive shaft. The gearbox has a gear shaft rotatable about a gear shaft axis. The gear shaft has a first end drivingly engaged to a first accessory and a second end drivingly engaged to a second accessory.

A gas turbine engine component comprises an airfoil with a suction side and pressure side extending from a leading edge to a trailing edge. There are a plurality of cooling holes adjacent the leading edge, with the cooling holes having a non-circular shape, with a longer dimension and a smaller dimension. The airfoil defines a radial direction from a radially outer end to a radially inner end, and radially outer of the cooling holes spaced toward the radially outer end, which have the longer dimension extending closer to parallel to the radial direction. Radially inner cooling holes closer to the radially inner end having the longer dimension extend to be closer to perpendicular relative to the radial direction compared to the radially outer cooling holes.

A seal assembly includes a first feather seal with a first cooling hole extending through the first feather seal. The seal assembly also includes a second feather seal adjacent to the first feather seal. The second feather seal includes a second cooling hole extending through the second feather seal. The first cooling hole is positioned over at least a portion of the second cooling hole.

An accessory system for a gas turbine engine having a driveshaft with an axis of rotation is provided. The system includes a towershaft coupled to the driveshaft and rotatable about a towershaft axis of rotation. The towershaft includes a towershaft bevel gear. The system includes a primary shaft including a first bevel gear and a second bevel gear that each revolve about a primary shaft axis of rotation. The first bevel gear is coupled to the towershaft bevel gear. The system includes a secondary shaft including a third bevel gear and a fourth bevel gear that each revolve about a secondary shaft axis of rotation. The third bevel gear is coupled to the second bevel gear. The system includes a tertiary shaft including a fifth bevel gear that revolves about a tertiary shaft axis of rotation. The fifth bevel gear is coupled to the fourth bevel gear.

An acoustic damper includes a low porosity layer section and a housing. The low porosity layer section is formed in a liner of a gas turbine combustor and has an arrangement of elongated generally S-shaped slots formed therein. The housing has a plurality of feed apertures. The housing is coupled to the low porosity layer section thereby defining a cavity such that air outside the housing is configured to flow through the apertures and through the elongated generally S-shaped slots in the low porosity layer section, thereby transforming acoustic energy into thermal energy and aiding in providing an acoustic dampening effect for the gas turbine combustor during operation thereof.

A method for designing and building a wheel that is simultaneously a turbine and an impeller with a plurality of impeller blades, wherein each impeller blade of the plurality of impeller blades is hollow along an entire length of the impeller blade and which leads into a peripheral circular chamber that operates as a fueled engine (THRA). The method includes building an evolving section of an inner channel of the impeller blades with a plurality of strips each having a neutral axis, wherein each impeller blade rests on a profile of a plurality of profiles of a corresponding neutral axis, the profile built for an inlet to the turbine and for inlets to each impeller blade.

A reverse flow gas turbine engine has a low pressure (LP) spool and a high pressure (HP) spool arranged sequentially in an axial direction. The LP spool comprises an LP compressor disposed forward of an LP turbine and drivingly connected thereto via an LP compressor gear train. The HP spool comprises an HP compressor in flow communication with the LP compressor, and an HP turbine disposed forward of the HP compressor and drivingly connected thereto via an HP shaft.