Airfoils and methods of cooling an airfoil are provided. The airfoil may comprise a spar; a coversheet on the spar; and a dual feed circuit between the spar and the coversheet. The dual feed circuit may include a first dam, a second dam spaced apart from the first dam along the chord axis of the spar, a first inlet disposed adjacent to the first dam, a second inlet disposed adjacent to the second dam, a circuit outlet disposed between the first inlet and the second inlet, and a plurality of diamond and/or hexagonal pedestals disposed on an outer surface of the spar. The diamond and/or hexagonal pedestals may form a plurality of cooling channels between the first inlet, the second inlet, and the circuit outlet. There may be no other circuit inlets are located between the first inlet and the second inlet.

A core structure for a providing a cooling passage in a gas turbine engine includes a core body that has a first cooling passage core. The first cooling passage core has a first width in a chord-wise direction near a first wall. A second width in the chord-wise direction near a second wall. A third width in the chord-wise direction between the first and second walls. The third width being smaller than the first and second widths to form an hourglass shape.

A dust mitigation system for airfoils includes a plurality of contoured tip turns which curve about at least two axes. This inhibits recirculation areas common within airfoils and further inhibits dust build up within the cooling flow path of the airfoil.

A composite layer system is presented. The composite layer system includes a metallic substrate, a structured surface, and a thermal protection system. The structured surface may be additively manufactured onto the metallic substrate and includes structured surface features formed to project above the metallic substrate. Each of the structured surface features are separated from adjacent structured surface features by grooves. The thermal protection coating may be thermally sprayed onto the structured surface and is bonded to each of the structured surface features.

A gas turbine system (1) including a burner arrangement having a tubular combustion chamber (5), a turbine (6) and a transition duct (7) connecting the combustion chamber (5) and the turbine (6), wherein the transition duct (7) is provided with an axially extending cooling air channel (11). The transition duct (7) includes a plurality of axially extending cooling air channels, and wherein each cooling air channel (11) is provided with one single inlet (12) opened to the outside of the transition duct (7) and with one single outlet (12) opened to the inside of the transition duct (7).

A ceramic matrix composite (CMC) component and method of fabrication including one or more elongate functional features in the CMC component. The CMC component includes a plurality of longitudinally extending ceramic matrix composite plies in a stacked configuration. Each of the one or more elongate functional features includes an inlet in fluid communication with a source of a cooling fluid flow. The CMC component further includes one or more bores cutting through the plurality of longitudinally extending ceramic matrix composite plies from at least one of the one or more elongate functional features to an outlet proximate to an outer surface of the ceramic matrix composite to form a cooling channel. The component may optionally include one or more film cooling throughholes cutting through the plurality of longitudinally extending ceramic matrix composite plies from an inner surface of the ceramic matrix composite component to an outlet proximate to the outer surface of the ceramic matrix composite component.

A turbine component includes a root and an airfoil extending from the root to a tip opposite the root. The airfoil forms a leading edge and a trailing edge portion extending to a trailing edge. A plurality of axial cooling channels in the trailing edge portion of the airfoil are arranged to permit axial flow of a cooling fluid from an interior of the turbine component at the trailing edge portion to an exterior of the turbine component at the trailing edge portion. A method of making a turbine component includes forming an airfoil having a trailing edge portion with axial cooling channels. The axial cooling channels are arranged to permit axial flow of a cooling fluid from an interior to an exterior of the turbine component at the trailing edge portion. A method of cooling a turbine component is also disclosed.

A turbine component includes a root and an airfoil extending from the root to a tip opposite the root. The airfoil forms a leading edge and a trailing edge portion extending to a trailing edge. Radial cooling channels in the trailing edge portion of the airfoil permit radial flow of a cooling fluid through the trailing edge portion. Each radial cooling channel has a first end at a lower surface at a root edge of the trailing edge portion or at an upper surface at a tip edge of the trailing edge portion and a second end opposite the first end at the lower surface or the upper surface. A method of making a turbine component and a method of cooling a turbine component are also disclosed.

A turbine component includes a root and an airfoil extending from the root to a tip opposite the root. The airfoil forms a leading edge and a trailing edge portion extending to a trailing edge. A plurality of nested cooling channels in the trailing edge portion of the airfoil permit passage of a cooling fluid from an interior of the turbine component to an exterior of the turbine component at the trailing edge portion. A method of making a turbine component includes forming an airfoil having a leading edge, a trailing edge portion extending to a trailing edge, and a plurality of nested cooling channels in the trailing edge portion. Each nested cooling channel fluidly connects an interior of the turbine component with an exterior of the turbine component at the trailing edge portion. A method of cooling a turbine component is also disclosed.

An assembly for a gas turbine engine according to an aspect of the present disclosure includes a metallic damper including a first contact surface and a gas turbine engine component. The gas turbine engine component includes a main body extending in a first direction between a gaspath surface and a second contact surface. The first and second contact surfaces oppose each other along an interface extending in a second direction. The first and second contact surfaces are dimensioned to contact each other along the interface in a hot assembly state. The main body is established by a composite including fibers in a matrix material. At least some of the fibers are arranged to establish a plurality of cooling passages aligned with the interface relative to the second direction. A method of damping for a gas turbine engine is also disclosed.