Core assemblies for manufacturing airfoils and airfoils for gas turbine engines are described. The core assemblies include a tip flag cavity core having an upstream portion, a tapering portion, and a downstream portion, with the tapering portion located between the upstream portion and the downstream portion and the downstream portion defines an exit in a formed airfoil. The upstream portion has a first radial height H.sub.1, the downstream portion has a second radial height H.sub.2 that is less than the first radial height H.sub.1, the tapering portion transitions from the first radial height H.sub.1 at an upstream end to the second radial height H.sub.2 at a downstream end, and at least one metering pedestal aperture is located within the tapering portion.

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.

The present invention relates to a guide vane cascade for a turbomachine, which has guide vanes that are mounted adjustably about an axis of rotation, so as to change an inflow angle each time, wherein a first guide vane and a second guide vane arranged on the pressure side thereof, referred to a longitudinal axis of the turbomachine, are arranged with an axial offset, namely, the axis of rotation of the second guide vane is offset axially toward the back, wherein the first guide vane and the second guide vane are provided such that, in an adjusted state, they form together a tandem configuration in a radially outer region but together they delimit a divergent channel in a radially inner region.

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 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.

A modular fluid-driven linear actuator comprising an actuator support comprising a first part and a second part configured for relative linear movement, wherein the actuator support is configured to hold a plurality of removable actuator modules in parallel and in respective module-receiving locations defined between the first part and the second part. At least one actuator module is removably installed in the actuator support and held in a respective module-receiving location of the plurality. A first fluid manifold is configured to provide a drive fluid to each actuator module when received in a respective module-receiving location, to act on a piston of the actuator module having a respective piston area to drive the piston in a first direction. A total piston area of the modular actuator is variable by installing and uninstalling actuator modules in the actuator support.

Methods of manufacturing or repairing a turbine blade or vane are described. The airfoil portions of these turbine components are typically manufactured by casting in a ceramic mold, and a surface made up of the cast airfoil and at the least the ceramic core serves as a build surface for a subsequent process of additively manufacturing the tip portions. The build surface is created by removing a top portion of the airfoil and the core, or by placing an ultra-thin shim on top of the airfoil and the core. The overhang projected by the shim is subsequently removed. These methods are not limited to turbine engine applications, but can be applied to any metallic object that can benefit from casting and additive manufacturing processes. The present disclosure also relates to finished and intermediate products prepared by these methods.

A component for a gas turbine engine includes a body portion that extends between a leading edge and a trailing edge of the component. The trailing edge includes a flared region and a non-flared region. At least one discharge slot is disposed at least partially within the flared region of the component.

Turbine distributor sector for an aircraft turbine engine, including an external annular platform sector and an internal annular platform sector, the sectors being coaxial and being connected together by blade assemblies including inner cavities cooled by gas circulation, the external platform sector including through openings of which radially internal ends open into the inner cavities, wherein the external platform sector includes inner ducts for supplying the cavities with gas, the ducts including air outlets opening into the openings and air inlets opening onto a portion of the external annular surface of the external platform sector.

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.