Abstract: An apparatus for an engine component in a turbine engine. The engine component including a wall with a cooling hole having a passage extending between an inlet fluidly coupled to a cooling fluid flow and an outlet at a heated surface. The cooling hole including a layup surface defining a first angle (?) and a layback surface defining a second angle (?).

Abstract: A component assembly for a gas turbine engine defining a core air flowpath is provided. The component assembly includes an outer shell comprising a first array of integral outer shell airfoils that extend inward from an outer shell periphery; and an inner shell comprising a second array of integral inner shell airfoils that extend outward from an inner shell periphery, wherein the outer shell and the inner shell are one or both of translatable and rotatable relative to one another between a first position and a second position.

Abstract: A component assembly for a gas turbine engine defining a core air flowpath is provided. The component assembly includes an outer shell comprising a first array of integral outer shell airfoils that extend inward from an outer shell periphery; and an inner shell comprising a second array of integral inner shell airfoils that extend outward from an inner shell periphery, wherein the outer shell and the inner shell are one or both of translatable and rotatable relative to one another between a first position and a second position.

Abstract: A component assembly for a gas turbine engine having a combustor defining a combustion chamber, the gas turbine engine defining a core air flowpath, a radial direction, and a circumferential direction is provided. The component assembly includes an outer shell at least partially defining the core air flowpath, the outer shell having an outer shell periphery that comprises a first array of integral outer shell airfoils that extend inward from the outer shell periphery; and an inner shell at least partially defining the core air flowpath, the inner shell having an inner shell periphery that comprises a second array of integral inner shell airfoils that extend outward from the inner shell periphery.

Abstract: A rotor assembly for a gas turbine engine includes a shaft and first and second annular drum segments coupled to the shaft. Furthermore, the rotor assembly includes an annular flange positioned between the first and second annular drum segments along the axial centerline, with the annular flange coupled to the first and second annular outer drum segments. Additionally, the rotor assembly includes a blade having a shank section and an airfoil section. The shank section is, in turn, coupled to the annular flange such the airfoil section extends inward along the radial direction toward the axial centerline and into a hot gas path of the gas turbine engine.

Abstract: A shroud assembly for a gas turbine engine including a plurality of rotor blades. The shroud assembly includes a plurality of tip shrouds, a plurality of flanges, and a plurality of first compressible elements. Each of the plurality of tip shrouds includes an outer band. Further each of the plurality of tip shrouds is coupled to one of the plurality of rotor blades at a tip end. Each of the plurality of flanges extends radially outward from one of the plurality of tip shrouds. Each of the plurality of first compressible elements is coupled to at least one of the plurality of flanges or one of the plurality of tip shrouds and oriented in a first circumferential direction. Further, each of the plurality of first compressible elements is oriented toward an adjacent tip shroud such that the tip shrouds mechanically engage to form a circumferential shroud.

Abstract: Shrouds and shroud segments for gas turbine engines are provided. In one exemplary aspect, a gas turbine engine includes a first rotor blade stage and a second rotor blade stage. The gas turbine engine also includes a shroud formed from a plurality of shroud segments that each have a body and a vane extending from the body. The vane is disposed within the flow path between the first rotor blade stage and the second rotor blade stage to form at least a portion of a nozzle stage. Further, the body of the shroud segment defines an outer wall of the flow path and spans from the first rotor blade stage to the second rotor blade stage. An impingement baffle segment couples the shroud segment with a casing of the gas turbine engine.

Abstract: A shroud assembly for a gas turbine engine including a plurality of rotor blades. The shroud assembly includes a plurality of tip shrouds, a plurality of flanges, and a plurality of first compressible elements. Each of the plurality of tip shrouds includes an outer band. Further each of the plurality of tip shrouds is coupled to one of the plurality of rotor blades at a tip end. Each of the plurality of flanges extends radially outward from one of the plurality of tip shrouds. Each of the plurality of first compressible elements is coupled to at least one of the plurality of flanges or one of the plurality of tip shrouds and oriented in a first circumferential direction. Further, each of the plurality of first compressible elements is oriented toward an adjacent tip shroud such that the tip shrouds mechanically engage to form a circumferential shroud.

Abstract: Shrouds and shroud segments for gas turbine engines are provided. In one exemplary aspect, a gas turbine engine includes a first rotor blade stage and a second rotor blade stage. The gas turbine engine also includes a shroud formed from a plurality of shroud segments that each have a body and a vane extending from the body. The vane is disposed within the flow path between the first rotor blade stage and the second rotor blade stage to form at least a portion of a nozzle stage. Further, the body of the shroud segment defines an outer wall of the flow path and spans from the first rotor blade stage to the second rotor blade stage. An impingement baffle segment couples the shroud segment with a casing of the gas turbine engine.