A single-crystal turbine blade and a method of making such single-crystal turbine blade are disclosed. During manufacturing, a secondary crystallographic orientation of the material of the single-crystal turbine blade is controlled based on a parameter of a root fillet between an airfoil of the single-crystal turbine blade and a platform of the single-crystal turbine blade. The parameter can be a location of peak stress in the root fillet expected during use of the turbine blade.

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a fan drive turbine and a high pressure turbine. The fan drive turbine drives the low pressure compressor and a gear arrangement to drive the fan section. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.

A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section, a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a low pressure turbine and a high pressure turbine. The low pressure turbine drives the low pressure compressor and the gear arrangement to drive the propulsor. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.

A single-crystal turbine blade and a method of making such single-crystal turbine blade are disclosed. During manufacturing, a secondary crystallographic orientation of the material of the single-crystal turbine blade is controlled based on a parameter of a root fillet between an airfoil of the single-crystal turbine blade and a platform of the single-crystal turbine blade. The parameter can be a location of peak stress in the root fillet expected during use of the turbine blade.

It is an objective of the invention to provide a Ni-based superalloy cast article exhibiting the corrosion resistance compatible or superior to the conventional ordinary precision cast articles and reducing the cost than the conventional ones while maintaining the mechanical properties compatible to the conventional ones. There is provided a Ni-based superalloy cast article including: 12.1 to 16 mass % of Cr; 4 to 16 mass % of Co; 3 to 5 mass % of Al; 2.1 to 3.3 mass % of Ti; 3.5 to 9 mass % of W; 1 to 2.4 mass % of Mo; 1.2 mass % or less of Nb; 0.005 to 0.05 mass % of B; 0.03 to 0.2 mass % of C; more than 0 mass % and 0.005 mass % or less of O; and the balance being Ni and impurities.

A method of repairing a superalloy component includes subjecting the superalloy component, including a repair area, to a phase agglomeration cycle, which includes stepped heating and controlled cooling of the component. The method further includes applying weld material to the repair area to create a weld surface; and covering the weld surface with brazing material. The component is then subjected to a braze cycle to produce a brazed component. The brazed component is cleaned, and the cleaned component is subjected to a restorative heat treatment to restore the microcrystalline structure and mechanical properties of the component.

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section including a fan having a plurality of fan blades, and a nacelle surrounding the plurality of fan blades, a compressor section including a low pressure compressor and a high pressure compressor, the low pressure compressor including a plurality of stages, and the high pressure compressor including 6 or more stages. A turbine section includes a fan drive turbine that drives the fan section through a gear arrangement, and including a second turbine that drives the high pressure compressor. A power ratio is provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor, the power ratio defined by the second power input divided by the first power input, and the power ratio is less than or equal to 1.0.

A rotor blade for a gas turbine, in particular an aircraft gas turbine, having a blade root element and a stream deflection portion adjoining the blade root element (12) in the longitudinal direction of the blade (RR); respective centroids (24) of blade cross-sectional areas of the stream deflection portion residing on a common stacking axis (26). It is provided that, starting from a first centroid (24) of a first blade cross-sectional area adjoining the blade root element (12), the stacking axis (26) extend within a cone (28) whose apex resides within the first centroid (24), and whose cone height (KH) extends orthogonally to the plane of the blade cross-sectional area; the angle (β) of the cone (28) being greater than 0° and smaller than or equal to 4°; preferably greater than or equal to 0.5° and smaller than or equal to 2°.

In a method for manufacturing a turbine engine element such as a blade or vane, the element has an airfoil. The method includes: applying a load across an assembly of a first cast portion of the airfoil and a second cast portion of the airfoil; and applying current across a junction of the first cast portion and the second cast portion to fuse the second cast portion to the first cast portion.

A method of manufacturing a wing element that is provided inside a gas turbine and through which a fluid passes and a method of manufacturing a blade are provided. The method of manufacturing the wing element includes preforming the wing element; disposing the wing element inside a mold; sequentially melting the wing element inside the mold along one direction using a heating device; and solidifying the melted wing element using a cooling device.