A turbine component includes a root and an airfoil extending from the root to a tip opposite the root. The airfoil forms a leading edge and a trailing edge portion extending to a trailing edge. Radial cooling channels in the trailing edge portion of the airfoil permit radial flow of a cooling fluid through the trailing edge portion. Each radial cooling channel has a first end at a lower surface at a root edge of the trailing edge portion or at an upper surface at a tip edge of the trailing edge portion and a second end opposite the first end at the lower surface or the upper surface. A method of making a turbine component and a method of cooling a turbine component are also disclosed.

A turbine component includes a root and an airfoil extending from the root to a tip opposite the root. The airfoil forms a leading edge and a trailing edge portion extending to a trailing edge. A plurality of nested cooling channels in the trailing edge portion of the airfoil permit passage of a cooling fluid from an interior of the turbine component to an exterior of the turbine component at the trailing edge portion. A method of making a turbine component includes forming an airfoil having a leading edge, a trailing edge portion extending to a trailing edge, and a plurality of nested cooling channels in the trailing edge portion. Each nested cooling channel fluidly connects an interior of the turbine component with an exterior of the turbine component at the trailing edge portion. A method of cooling a turbine component is also disclosed.

An aerofoil component is formed of continuous fibre-reinforced polymer composite created by curing laid up pre-preg layers extending in radial and chordal directions of the aerofoil component, and further includes a plurality of Z-pins arranged in a pattern forming a chevron on the pressure and/or suction surface of the aerofoil component, the chevron having a vertex and two arms extending at an angle from each other away from the vertex either towards the radially inner root of the aerofoil component or towards the radially outer tip of the aerofoil component.

A method for joining two structural elements by welding, in particular by butt welding comprises forming a weld line joining the two structural elements; and adding material across the weld line, thereby forming one or more crack stoppers for limiting crack propagation along the weld line. The one or more crack stoppers each have a limited extension along the weld line as seen in relation to a length of the weld line. A structural system comprising two structural elements joined by the method is disclosed. The method may be applied, e.g., to components of aircraft engines.

Airfoil assemblies for gas turbine engines are described. The airfoil assemblies include an airfoil body having a leading edge, a trailing edge, a pressure side, and a suction side, the airfoil body extending in a radial direction between a first end and a second end, wherein the airfoil defines an internal cavity bounded by interior surfaces of the airfoil body, the airfoil body formed from a high-temperature-material material and a metallic insert member installed within the internal cavity. One or more radially extending ribs are arranged on an exterior surface of the metallic insert member and defining one or more radially extending passages between the exterior surface of the metallic insert member and the interior surface of the airfoil body.

An effusion cooling element for the surface of a hot gas path (HGP) component is disclosed. The effusion cooling element includes a coolant swirling chamber embedded within the body of the HGP component. A coolant delivery passage is in the body and configured to deliver a coolant to the coolant swirling chamber. The coolant swirling chamber imparts a centrifugal force to the coolant. An effusion opening is in the HGP surface and in fluid communication with the coolant swirling chamber, the effusion opening having a smaller width than the coolant swirling chamber. The coolant exits the effusion opening over substantially all of 360° about the effusion opening, creating a coolant film on the HGP surface.

In one exemplary embodiment, a gas turbine engine is provided. The gas turbine engine defines a radial direction, an axial direction, and an axis extending along the axial direction of the gas. The gas turbine engine includes: a shaft configured to rotate about the axis; an electric machine comprising a rotor coupled to and rotatable with the shaft and a stator, the rotor defining an end along the axial direction; and a cooling manifold rotatable with the rotor and positioned at the end of the rotor, the cooling manifold configured to receive a flow of cooling fluid and provide the cooling fluid to the stator during operation of the gas turbine engine.

The present invention relates to a turbine engine guide (23) vane (25), with a height (H) extending between a vane root (26) and a vane tip (27) along a radial direction (Z), said vane (25) comprising a succession of five bulge portions along a tangential direction (Y) perpendicular to the radial direction (Z), this succession of bulge portions extending over the whole height (H) of the vane (25), and the convexity of the successive bulge portions being alternately in one direction and in the other. The vane (25) has the advantage of having an eigenfrequency for the first striped vibration mode which is different from the urging frequencies of said vane (25), during the operation of the turbine engine.

An aeroplane turbojet low-pressure compressor vane includes a leading edge, a trailing edge, a surface, and an extrados surface which extend from the leading edge to the trailing edge. To combat the presence and the formation of ice, the vane is provided with an electric de-icing device with a thermistor. The thermistor forms a heating electrical track suitable for de-icing the vane. The present application also proposes a method for producing a turbine engine vane.

A profiled structure for an aircraft or turbomachine is elongated in a direction in which the structure has a length exposed to an airflow and includes serrations defined by successive teeth and depressions. The serrations may be transverse to a leading edge and/or a trailing edge of the profiled structure and in the direction of elongation. Along the profiled leading edge and/or profiled trailing edge, the successive teeth and depressions may extend only over a part of the length exposed to the flow. The amplitude and/or spacing of the teeth may vary monotonically except for the few teeth nearest each end of the part, with a remaining part of the length being smooth.