An apparatus and method an engine component for a turbine engine comprising an outer wall bounding an interior and defining a pressure side and an opposing suction side, with both sides extending between a leading edge and a trailing edge to define a chord-wise direction, and extending between a root and a tip to define a span-wise direction, at least one cooling passage located within the interior, at least one cooling hole having an inlet fluidly coupled to the cooling passage and an outlet located along the outer wall.

The present disclosure generally relates to variable cellular structures, methods of making such cellular structures, and variable cellular flow discouragers for turbine engines for jet aircraft.

An air turbine starter having a housing, a turbine, a drive shaft, and at least one vane. The housing having an inlet, an outlet and a curvilinear flow path extending between the inlet and the outlet. The at last one vane is located within a portion of the curvilinear flow path, and includes an outer wall extending between a root and a tip in a span-wise direction and between a leading edge and a trailing edge in a chord-wise direction. In one aspect, the vane is arranged to define an acute axial angle that is non-constant in the chord-wise or span-wise direction. In another aspect, the vane is arranged to define an acute tangential angle that is non-constant in the chord-wise or span-wise direction.

A modified J-type cantilevered vane reduces rubbing of an airfoil and enhances the vibration stability of a vane hub and the fluidity of fluid. The cantilevered vane includes a root attachment inserted into a slot formed in an inner circumferential surface of a casing of a gas turbine, and an airfoil vertically extending from the root attachment to a predetermined height. The airfoil includes a linear part extending in a radial direction from the inner circumferential surface of the casing toward a rotating shaft of the gas turbine, and a curved part integrally formed with the linear part, the curved part including a bend beginning from one end of the linear part and curving in a circumferential direction of the rotating shaft. Overall, the linear part and the curved part are biased in the circumferential direction, since the airfoil has a J-shaped X-axis profile and a C-shaped Y-axis profile.

Turn caps for airfoils of gas turbine engines having a first pressure-side turn passage extending from a respective inlet to a respective outlet within the turn cap, a first suction-side turn passage extending from a respective inlet to a respective outlet within the turn cap, and a merging chamber fluidly connected to the outlets of the first pressure-side turn passage and the first suction-side turn passage, wherein each of the first suction-side turn passage and the first pressure-side turn passage turn a direction of fluid flow from a first direction to a second direction such that a fluid flow exiting the first suction-side turn passage and the first pressure-side turn passage are aligned when entering the merging chamber.

A cooling circuit for a multi-wall blade according to an embodiment includes: a pressure side cavity with a surface adjacent a pressure side of the multi-wall blade; a suction side cavity with a surface adjacent a suction side of the multi-wall blade; a first leading edge cavity with surfaces adjacent the pressure and suction sides of the multi-wall blade, the first leading edge cavity located forward of the pressure and suction side cavities; and a second leading edge cavity with surfaces adjacent the pressure and suction sides of the multi-wall blade, the second leading edge cavity located forward of the first leading edge cavity.

An assembly for a turbine engine or turbine engine test bed, including a hub defined about a longitudinal axis, a blade, including a blade root, extending in a radial direction to the longitudinal axis, the hub including a recess capable of receiving the blade root by insertion in the radial direction, wherein the recess includes a groove, the blade root including a groove, so that the two grooves are facing each other when the blade root is inserted into the recess, the assembly further includes a seal located in the two grooves.

A gas turbine engine has a stator vane assembly. The stator vane assembly includes an inner diameter shroud, an outer diameter shroud located radially outward from the inner diameter shroud, a vane extending radially outward from the first inner diameter shroud to the outer diameter shroud. The wedge clip is positioned horizontally through the vane to prevent the vane from being dislodged from the stator vane assembly.

A component for a turbine engine comprises a wall with a surface along which a hot airflow passes, a second surface along which a cooling airflow passes, and an insert mounted to the wall wherein the material used for the insert can have a higher temperature capability than that of the wall.

A turbine vane of a turbine engine such as a turbojet. The vane includes a base supporting a blade that extends in a spanwise direction and ends in an apex. The blade includes a pressure-side wall and a suction-side wall, each ending at an end edge at the apex of the blade. The blade includes, at the apex thereof, a closed wall extending from the pressure-side wall to the suction-side wall so as to define, with the end edges, a tub shape. The pressure-side wall is curved inward so as to deviate from the spanwise direction in the region of the blade apex preceding the tub, between the base and the pressure-side end edge.