A turbine stator blade made of ceramic matrix composite material includes a hollow blade profile and has a trailing edge and a leading edge, the blade including a first portion including an extrados face and a second portion distinct from the first portion including an intrados face, the first and second portions being connected to one another by a connection interface present at least on the trailing edge or leading edge, the connecting interface including a region of overlap between the first and second portions present on at least one longitudinal end of the blade profile and intended to be present outside a flow path of a gas stream of the turbine, the blade also including a platform present at a longitudinal end of the blade profile and that includes a first portion integral with the extrados face and a second part integral with the intrados face.

A turbine blade tip, turbine blade and method where improved cooling is made possible by an improved cooling structure with cooling air holes inside a depression in a blade tip and a special arrangement of multiple cooling air holes which are supplied by a single cooling air channel inside a wall.

An airfoil for a turbine engine having a working airflow separated into a cooling airflow and a combustion airflow, the airfoil comprising a wall defining an interior and having an outer surface over which flows the combustion airflow, the outer surface defining a first side and a second side extending between a leading edge and a trailing edge to define a chord-wise direction; at least one cooling conduit located within the interior and fluidly coupled to the cooling airflow. A primary cooling passage having at least one inlet fluidly coupled to the at least one cooling conduit, a primary outlet on the outer surface. A passage connecting the at least one inlet to the primary outlet, the passage separated into a first portion and a second portion. The primary outlet spaced from the trailing edge a predetermined distance.

In one exemplary embodiment of the present disclosure an airfoil is provided. The airfoil defines a spanwise direction, a root end, and a tip end. The airfoil includes: a body defining a pressure side and a suction side and extending along the spanwise direction between the root end and the tip end, the body formed of a composite material; and a spar enclosed in the body of the airfoil extending along the spanwise direction, the spar comprising a plurality of spar segments arranged in an overlapping configuration along the spanwise direction.

This composite material blade, which is formed using a composite material including reinforcing fibers and resin, and which has a positive pressure surface and a negative pressure surface, is provided with a ventral part, being the part on the positive pressure surface side in a blade thickness direction, which is the direction joining the positive pressure surface and the negative pressure surface, a dorsal part, being the part on the negative pressure surface side in the blade thickness direction, and a metal shield portion which is provided on the leading edge side, being the upstream side in a flow direction in which a fluid flows, wherein: the metal shield portion includes a main body portion provided on the leading edge side, and an embedded portion which is provided on the trailing edge side, being the downstream side in the flow direction, of the main body portion, and which is provided between the ventral part and the dorsal part; and the plate thickness of the metal shield portion in the blade thickness direction decreases from the main body portion toward the embedded portion.

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.

Airfoils and core assemblies for such airfoils are described. The airfoils include a leading edge cavity defined, in part, by a leading edge interior rib and a trailing edge cavity defined, in part, by a trailing edge interior rib. A plurality of pressure side cavities are defined by pressure side skin cavity walls with at least one pressure side skin cavity wall not extending to the suction side wall. A plurality of suction side cavities are defined by suction side skin cavity walls with at least one suction side skin cavity wall not extending to the pressure side wall. A main body cavity extends between the leading edge interior rib and the trailing edge interior rib and the plurality of side cavities are arranged in a staggered pattern to define the bounds of the main body cavity.

An airfoil for a fan section of a turbine engine may include a fan blade or an outlet guide vane formed of a first material, and an edge guard disposed about an edge of the fan blade. The edge guard may include a matrix composite that has a toughness that is greater than a toughness of the first material. The airfoil may include a fan blade or an outlet guide vane. The first material of the airfoil may include a metal alloy and/or a matrix composite. A method of manufacturing an airfoil for a fan section of a turbine engine may include manufacturing an edge guard, attaching the edge guard to the airfoil.

An airfoil for a fan section of a turbine engine may include a fan blade or an outlet guide vane, and an edge guard attached thereto. The edge guard may include a heating conduit disposed within at least a portion of the edge guard. An anti-icing system for a plurality of fan blades or outlet guide vanes may include a fluid supply pathway configured to supply heating fluid to respective ones of a plurality of heating conduits within the edge guards attached to respective ones of a plurality of fan blades and/or to a plurality of outlet guide vanes. The heating fluid may include bleed air from a core air flowpath. A method of inhibiting icing on an airfoil may include flowing a heating fluid into a heating conduit disposed within an edge guard attached to the airfoil and heating the edge guard with the heating fluid.

A turbine vane assembly adapted for use in a gas turbine engine includes a support and a turbine vane arranged around the support. The support is made of metallic materials. The turbine vane is made of ceramic matrix composite materials to insulate the metallic materials of the support.