A component for a gas turbine engine, comprising: first and second walls; a coolant channel defined by the space between the first and second walls; and a first rib extending between the first and second walls to the end of the coolant channel in a coolant flow direction, such that the coolant channel is bifurcated in the coolant flow direction.

An aircraft turbine engine blade includes at least one inner cavity for circulating a ventilation air flow and having a wall with first projecting elements oriented in a first direction and forming air flow disrupters, and at least a second projecting element oriented in a second direction different from the first direction. The second projecting element and at least one of the first projecting elements overlap each other in one area. At least one of the first projecting elements overlaps the second projecting element and has a height (H2, H4′) which is greater than that of the second projecting element in the area and greater than that of the other first projecting elements of the wall, in order to retain its disruptive function along the entire length thereof.

A turbine wheel (100) for a turbocharger includes a hub (110) and blades (120). The hub includes a primary outer surface (116), a back wall surface (115), a peripheral edge (119) extending between the primary outer surface and the back wall surface, and a central axis (111). The hub has a back wall thickness measured parallel to the central axis from the primary outer surface to the back wall surface. The blades (120) extend from the primary outer surface of the hub and are integrally formed with the hub by casting. Each blade has a blade thickness measured tangential to the central axis between opposed surfaces of the blade. A maximum diameter of the peripheral edge is 60 mm or more. Over a majority of a radially outer region (117), a thickness ratio of the blade thickness to the back wall thickness varies by 25% or less. The radially outer region extends from the peripheral edge to a 25% meridional.

A nozzle segment for a gas turbine engine comprises an outer band having a cooling air inlet, an inner band having a first cooling air outlet, and a nozzle airfoil comprising a cooling flow passage arranged to receive the cooling air as a cooling air stream. A first channel and a second channel are arranged within the cooling flow passage. A deflector is arranged to divide the cooling air stream into a first cooling air stream in the first channel and a second cooling air stream in the second channel, respectively. The deflector deflects the first cooling air stream obliquely to a suction sidewall in the first channel, wherein the first channel is configured to transport the first cooling air stream along the first channel in a swirly flow.

A gas turbine engine article includes a cooling passage network embedded in an article wall between inner and outer portions of the article wall. The network has an inlet orifice through the inner portion to receive cooling air from a cavity, a sub-passage region that includes an array of pedestals, and at least one outlet orifice through the outer portion. The array of pedestals includes first pedestals arranged in a first row and second pedestals arranged in a second, adjacent row. The first pedestals and the second pedestals define inter-row sub-passages there between. Each of the inter-row sub-passages has an inlet mouth, an outlet mouth, and a compound channel connecting the inlet mouth and the outlet mouth. The compound channel includes a first channel length over which the inter-row sub-passage has a constant cross-section and a second channel length over which the inter-row sub-passage has a non-constant cross-section.

A system for automated adaptive manufacturing of airfoil castings is disclosed. The system may receive a three dimensional scan of a work piece. The system may compare the three dimensional scan to a digital model of the work piece. The system may identify an area of dimensional abnormality on the work piece based on the comparison. The system may alter the area of dimensional abnormality on the work piece.

Core assemblies for manufacturing airfoils and airfoils made therefrom are described. The core assemblies having a leading edge hybrid skin core positioned relative to a plurality of core bodies and configured to define a leading edge cavity at a leading edge of the manufactured airfoil. The leading edge hybrid skin core extends from a root region toward a tip region in a radial direction, the leading edge hybrid skin core extends above at least one of the plurality of core bodies to define an exit in a tip region of the manufactured airfoil, and the leading edge hybrid skin core has a height-to-width ratio of about 0.8 or less.

A method of manufacturing for finishing ceramic core flash. Locating a first hole on a ceramic core by a laser sensor on a robot. Probing for a center of the first hole by a force-torque sensor on the robot. Scanning for an axial position of a second hole on the ceramic core. Scanning for a third hole on the ceramic core. Probing for a center of the third hole. Determining axial and radial scale factors based on the first hole location and the third hole location. Uploading the axial and radial scale factors to the robot Multiplying the X component position by the axial scale factor and the Z component position by the radial scale factor in an array format. Cutting a designated scaled location along the ceramic core to remove flash. Repeating process for additional scaled locations along the ceramic core.

A method of manufacturing a casting is provided and includes establishing desired dimensions of a nominal casting, executing a casting process to produce multiple actual castings with each of the multiple actual castings having respective dimensions that differ from each other and from the desired dimensions of the nominal casting and engaging one or more tools to adaptively machine, without rigidly-programmed toolpaths, each of the multiple actual castings to reduce the respective differences between the actual dimensions of each of the multiple actual castings and the desired dimensions.

Methods for manufacturing an impeller wheel assembly (e.g., an impeller wheel attached to one or more additional components) are provided. In one example, a method includes casting an impeller wheel without a hub feature in a mold and holding the cast impeller wheel in a fixed position using a holding plate during subsequent stages of fabrication (e.g., the addition of a shaft via friction welding).