A set of compression stage(s) of a turbomachine, forming an annular fluid passage and comprising at least one annular stator platform and/or at least one annular rotor platform having an outer longitudinal profile inclined (I.sub.Si/I.sub.Ri) inwards and upstream with respect to a nominal profile of the fluid stream, where the inclination (I.sub.Si/I.sub.Ri) of the outer longitudinal profile of the or each of the annular platforms, relative to the nominal profile of the fluid stream, oscillates along the circumference of the annular platform or platforms, between a maximum value in front of the blades of the annular platform and a minimum value between each pair of adjacent blades of the annular platform.

A vane arc segment includes an airfoil fairing that has first and second fairing platforms and an airfoil section therebetween. The airfoil section has a pressure side, a suction side, and an internal cavity. The first fairing platform defines suction and pressure side circumferential mate faces, forward and aft faces, a gaspath side, a non-gaspath side, and a flange that projects from the non-gaspath side. The flange extends along one of the suction or pressure side circumferential mate faces. There is a rib that has a rib section in the internal cavity that spans the pressure and suction sides and a rib extension section that extends from the rib section in the internal cavity and along the non-gaspath side of the first fairing platform. The rib extension intersects the flange to form a gusset for reinforcing the flange.

A turbine vane includes an airfoil having a leading edge and a trailing edge, an inner shroud disposed at one end of the airfoil to support the airfoil, and an outer shroud disposed opposite to the inner shroud at the other end of the airfoil to support the airfoil, wherein a corner part is formed at a portion where the airfoil and the inner shroud or the outer shroud meet, the corner part including a first round portion connected in an arc shape to the inner shroud or the outer shroud, a first inclined portion connected to the first round portion and outwardly extending in an inclined shape, and a second round portion connected to the first inclined portion and outwardly extending in an arc shape.

An airfoil including: an airfoil portion having an airfoil surface; and a communication hole extending at least in the airfoil portion and a first opening open in the airfoil surface, through which the first opening is communicated with a second opening provided in a portion of the airfoil. On a cross-section perpendicular to the spanwise direction through a position of the first opening of the spanwise direction, an angle A1 satisfying a condition (a) exists within an angle range −10 degrees to 10 degrees with respect to an extension line obtained by extending a camber line of the airfoil portion from a leading edge. The condition (a) is a static pressure at a position of the first opening is equal to a static pressure at a position of the second opening when the airfoil portion receives a fluid flow from a direction of the angle A1 toward the leading edge.

To increase the inertia of a sealing lip of a blade for an aircraft turbine engine, and thus improve the service life of such a sealing lip, the sealing lip is conformed so as to have a trough in the outer surface thereof and a corresponding boss in the inner surface thereof, the trough and the boss being defined based on a connection cross section of the sealing lip to a blade body, and being formed at a distance from a free axial end of the sealing lip.

A turbomachine defines an axial direction, a radial direction perpendicular to the axial direction, and a circumferential direction extending concentrically around the axial direction. The turbomachine includes a nozzle having an inner platform, an outer platform, and an airfoil. The airfoil includes a leading edge, a trailing edge downstream of the leading edge, a pressure side surface, and a suction side surface opposite the pressure side surface. The trailing edge is orthogonal with the outer platform in an axial-radial plane and the trailing edge is oblique to the inner platform in the axial-radial plane.

A gas turbine engine component has a component body configured to be positioned within a flow path of a gas turbine engine having an external pressure, and wherein the component body includes at least one internal cavity having an internal pressure. At least one inlet opening is formed in an outer surface of the component body to direct hot exhaust gas flow into the at least one internal cavity, and there is at least one outlet from the internal cavity. The internal pressure is less than an inlet external pressure at the inlet opening and the internal pressure is greater than an outlet external pressure at the outlet opening to controllably ingest hot exhaust gas via the inlet opening and expel the hot exhaust gas via the outlet opening to maintain a laminar boundary layer along the outer surface of the component body.

A method is provided involving a fairing for a turbine engine. During this method, a detail is provided. The detail includes a carrier and an exterior layer bonded to the carrier. The carrier is configured from or otherwise includes fiber-reinforced composite material. The exterior layer is configured from or otherwise includes polymer material. The detail is arranged with the fairing. The fairing includes an exterior side and an edge. The detail covers and extends along at least a portion of the exterior side. The detail wraps at least partially around the edge. The carrier and at least a first overhang portion of the exterior layer are bonded to the fairing.

A vane arc segment includes a continuous airfoil piece that defines first and second platforms and an airfoil section that extends between the first and second platforms. The airfoil section has a pressure side and a suction side. The first platform defines axially-sloped suction and pressure side circumferential mate faces, first and second axial sides, a gaspath side, a non-gaspath side, and a flange that projects from the non-gaspath side. The flange extends along at least a portion the axially-sloped suction side circumferential mate face.

A gas turbine engine includes a rotor rotatable about a central axis. The gas turbine engine includes a turbine stage including a stationary portion and a rotating portion made up of a number of rotating blades and a plurality of stationary vanes arranged to define the stationary portion. Each stationary vane includes an inner rail having an inlet face, a suction side face, a pressure side face, and a platform. A vane portion extends along a radial line from the platform and defines one of a first stagger angle and a second stagger angle with respect to the central axis. The platform has an elliptical cross-section in a plane that includes the central axis.