Abstract: A vane arc segment includes an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the flange. There is a conformal thermal insulation blanket disposed on the flange.

Abstract: A gas turbine engine component according to an example of the present disclosure includes a wall extending in a thickness direction between first and second wall surfaces. The first wall surface bounds an internal cavity. The wall includes a plurality of cooling passages. Each of the cooling passages extend in a first direction between an inlet port and an outlet port coupled to a respective diffuser, and the inlet port coupled to the internal cavity along the first wall surface. Sidewalls of adjacent diffusers are conjoined to establish a common diffuser region interconnecting the diffusers and a common outlet along the second wall surface. A method of cooling a gas turbine engine component is also disclosed.

Abstract: A gas turbine engine component includes a wall that has an inner surface and an outer surface. An inlet is defined by the inner surface. At least one non-rectangular slot is defined by the outer surface and includes at least one protrusion extending into the slot. A slot passage fluidly connects the inlet to the at least one non-rectangular slot. The slot passage comprises an inlet portion that extends through the wall from the inlet to an intermediate portion. An outlet portion extends through the wall from the intermediate portion to the at least one non-rectangular slot.

Abstract: An airfoil for a gas turbine engine is disclosed. In various embodiments, the airfoil includes a cooling passage; an outer wall separating a core flow path from the cooling passage; a diffuser in fluid communication with the cooling passage and opening into the core flow path, the diffuser being characterized by a linear ridge on a downstream end of the diffuser; and a thermal barrier coating covering the outer wall and the linear ridge.

Abstract: An airfoil vane includes an airfoil section including an outer wall that defines an internal cavity; and a baffle situated in the internal cavity, the baffle including a baffle wall that defines a central cavity having a leading end and a trailing end corresponding to a leading end and a trailing end of the airfoil section, and a tail extending from the baffle wall, the tail including at least one feature configured to disturb an airflow surrounding the tail. A baffle for the airfoil vane assembly and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.

Abstract: A component for a gas turbine engine includes a trailing edge tip corner that at least partially defines a trailing edge cavity and a multiple of corner features within the trailing edge cavity, the multiple of corner features splayed along the trailing edge tip corner.

Abstract: An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section that has an internal wall and an external wall. The external wall defines pressure and suction sides that extends in a chordwise direction between a leading edge and a trailing edge, a first impingement cavity and a second impingement cavity bounded by the external wall at a leading edge region that defines the leading edge. A first crossover passage within the internal wall is connected to the first impingement. The first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity. A second crossover passage within the internal wall is connected to the second impingement cavity. The second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity.

Abstract: A gas turbine engine component according to an example of the present disclosure includes a wall extending in a thickness direction between first and second wall surfaces. The first wall surface bounds an internal cavity. The wall includes a plurality of cooling passages. Each of the cooling passages extend in a first direction between an inlet port and an outlet port coupled to a respective diffuser, and the inlet port coupled to the internal cavity along the first wall surface. Sidewalls of adjacent diffusers are conjoined to establish a common diffuser region interconnecting the diffusers and a common outlet along the second wall surface. A method of cooling a gas turbine engine component is also disclosed.

Abstract: A gas turbine engine component according to an example of the present disclosure includes, among other things, an external wall extending in a thickness direction between first and second wall surfaces. The first wall surface bounds an internal cavity, and establishes at least one surface depression along an external surface contour. The external wall includes at least one cooling passage having an outlet port established along the at least one surface depression. A method of fabricating a gas turbine engine component is also disclosed.

Abstract: A component for a gas turbine engine includes, among other things, an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge, the airfoil extending radially from a platform to a tip, a tip pocket formed in the tip and terminating prior to the trailing edge, and one or more heat transfer augmentation devices formed in the tip pocket.

Abstract: An airfoil for a gas turbine engine is disclosed. In various embodiments, the airfoil includes a cooling passage; an outer wall separating a core flow path from the cooling passage; a diffuser in fluid communication with the cooling passage and opening into the core flow path, the diffuser being characterized by a linear ridge on a downstream end of the diffuser; and a thermal barrier coating covering the outer wall and the linear ridge.

Abstract: An airfoil includes a ceramic airfoil section that defines leading and trailing edges, pressure and suction sides, and radially inner and outer ends. The span of the airfoil section has first, second, and third radial span zones. There is a row of cooling holes in an aft 25% of the axial span. The row of cooling holes extends though the first, second, and third radial span zones. The cooling holes in the first radial span zone define a first pitch P1, the cooling holes in the second radial span zone define a second pitch P2, and the cooling holes in the third radial span zone define a third pitch P3, wherein P2<P3<P1.

Abstract: A combustor panel of a combustor may include a combustion facing surface, a cooling surface opposite the combustion facing surface, and heat transfer pins extending from the cooling surface. A grouping of the heat transfer pins may include a metallic coating.

Abstract: A component for a gas turbine engine includes, among other things, an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge, the airfoil extending radially from a platform to a tip, a tip pocket formed in the tip and terminating prior to the trailing edge, and one or more heat transfer augmentation devices formed in the tip pocket.

Abstract: A liner panel for use in a combustor of a gas turbine engine, the liner panel including a cold side; and a rail that extends from the cold side, the rail includes a first diffusion interface passage surface and a second diffusion interface passage surface, the first diffusion interface passage surface angled with respect to the second diffusion interface passage surface.

Abstract: An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges. A stagnation line is located on the pressure side wall aft of the leading edge. A cooling passage is provided between the pressure and suction side walls. Forward-facing cooling holes are provided adjacent to the stagnation line on the pressure side wall and oriented toward the leading edge.

Abstract: A gas turbine engine includes a compressor, a combustor fluidly connected to the compressor via a core flowpath, and a turbine fluidly connected to the combustor via the core flowpath. The turbine includes at least one stage having a plurality of rotors and a plurality of vanes. An outer diameter of the core flowpath at at least one stage is at least partially defined by a set of circumferentially arranged blade outer air seals. Each blade outer air seal includes a platform. An internal cooling cavity is defined within the platform. At least one mateface of the platform includes a cooling trench, and a first set of cooling holes connecting the internal cavity to the cooling trench.

Abstract: An airfoil includes a leading edge and a trailing edge. A first exterior wall extends between the leading edge and the trailing edge. A second exterior wall opposite the first exterior wall extends between the leading edge and the trailing edge. A first cavity is adjacent the first exterior wall and includes a central portion and at least one forward extending slot passage and at least one aft extending slot passage.

Abstract: A lost core mold component comprises a first leg and a second leg with a plurality of crossover members connecting the first and second legs. The plurality of crossover members includes outermost crossover members spaced from each other. Adjacent ends of each of the first and second legs, and second crossover members are spaced closer to each other than are the outermost crossover members. Central crossover members extend between the first and second leg and between the second crossover members. The outermost crossover members extend for a first cross-sectional area. The second crossover members extend for a second cross-sectional area and the central crossover members extend for a third cross-sectional area. The first cross-sectional area is greater than the second cross-sectional area. The second cross-sectional area is greater than the third cross-sectional area. A gas turbine engine and component are also disclosed.

Abstract: A combustor of a gas turbine engine including a first forward liner panel. The forward liner panel including a liner panel forward edge, a liner panel aft edge opposite of the liner panel forward edge, and a first liner panel side edge between the liner panel aft edge and the liner panel forward edge. The first liner panel side edge is non-perpendicular to the liner panel forward edge and oriented at an angle about perpendicular to a swirl flow in the combustor.