A method is provided of processing one or more blades of a row of blades which forms part of a rotor for a gas turbine engine. The method includes: providing a rotor disc having a slot for mounting one or more blades; applying tape at one or both sides of the slot such that when the blades are loaded in the slot the tape reduces a range of tilt angles which can be adopted by the blades; loading the blades in the slot; and performing a processing operation on the loaded blades.

A method for repairing a turbine housing of an air cycle machine is provided. The method comprises: removing damaged material from a journal bearing bore of the turbine housing; forming the journal bearing bore to an inner diameter of about 2.580±0.001 inches (6.553±0.003 centimeters). The journal bearing bore having an inlet and an outlet. The method also comprises determining the inner diameter of the journal bearing bore; and forming a cylindrical insert having a first end, a second end, an inner surface, a step located at the first end, and an outer surface with an outer diameter that provides a diametric interference with the inner diameter of the journal bearing bore. The diametric interference being between about 0.000-0.003 inches (0.000-0.008 centimeters). The method further comprises inserting the second end of the cylindrical insert into the inlet of the journal bearing bore up to the step.

A turbocharger includes a turbine housing and a wastegate assembly. The turbine housing has a wastegate duct defining a wastegate channel and a valve seat disposed about the wastegate channel. The wastegate assembly has a valve body engageable with the valve seat and a shaft fixed to the valve body at a connection interface for selectively engaging the valve body with the valve seat at a first plane. The connection interface has a centroid and a second plane tangential to the centroid. A second axis is normal to the second plane and is angularly tilted relative to a first axis normal to the first plane at which the valve body and the valve seat are engaged.

A blade outer air seal (BOAS) includes a body defining a core having a forward cavity portion and a plurality of forward hooks extending radially from the body, the hooks extending to a hook height. The BOAS includes a seal surface defined between the forward hooks and which includes a forward side, an aft side, and a seal surface height. The seal surface is aft of the forward cavity portion. The BOAS includes a thermal regulation channel defining an inlet aft of the sealing surface and an outlet defined in fluid communication with the forward cavity portion.

The present invention relates to a relief valve (1) for a turbocharger, in which the crank arm (3) is made of a first material and the shaft (4) is made of a second material different from the first material used for manufacturing the crank arm (3), each of the materials containing a composition that provides the necessary properties according to the application of each component of the relief valve (1). The present invention also relates to a process for manufacturing the relief valve (1), which allows the crank arm (3) and the shaft (4) to be manufactured separately, using different materials for the manufacture of each component.

A high-temperature component including a plurality of cooling passages through which the cooling medium can flow, a header connected to respective downstream ends of the plurality of cooling passages, and one or more outlet passages for discharging the cooling medium flowing into the header to outside of the header. The one or more outlet passages are less in number than the plurality of cooling passages. Respective minimum flow passage cross-sectional areas of the one or more outlet passages are not less than respective flow passage cross-sectional areas of the plurality of cooling passages in a connection between the header and the cooling passages. A sum of the respective minimum flow passage cross-sectional areas of the one or more outlet passages is less than a sum of the respective flow passage cross-sectional areas of the plurality of cooling passages in the connection between the header and the cooling passages.

The invention relates to a component for a thermal machine, in particular a gas turbine, which includes a corner and/or edge subjected to a thermally high load. The cooling of the component is improved in a manner such that at least one cooling channel is countersunk into the surface of the component in the immediate vicinity of the corner and/or edge in order to cool the corner and/or edge.

In an adjustable guide vane mechanism for a turbine, comprising a bearing ring with a plurality of guide vanes rotatably supported on the bearing ring by means of guide vane shafts provided with actuating levers which are engaged by a rotatable ring, the rotatable ring comprises axially spaced radial inner and outer sections with a radially inclined center section joining the spaced radial inner and outer sections, the radially inner section being rotatably supported and the radially outer section extending around the actuating levers and having recesses in which the actuating levers are received for pivoting the actuating levers and the guide vanes upon rotation of the rotatable ring.

The invention relates to an aircraft turboprop engine having two unducted propellers, coaxial about a rotation axis (LL), respectively upstream (12) and downstream (14′), each propeller comprising an annular row of blades, the blades of the downstream propeller (14′) each having a truncated head and each comprising a leading edge (20′) and a trailing edge (22′), the radially external ends (A2, B1) of which are distant from each other and are connected by a terminal edge (21′), characterised in that the terminal edge (21′) of each blade of the downstream propeller (14′) has, at at least two points, tangents in a meridian plane having different slopes, the upstream end (A) of the terminal edge defining a maximum radial dimension (R2) of the blade and its downstream end (B1) defining a minimum radial dimension (R3) of the terminal edge.

In an exemplary embodiment, a method for manufacturing turbine wheel airfoils includes: defining an initial design with an initial respective line for a straight line cut for a respective surface of each airfoil; evaluating an initial score for the initial design based on mechanical, aerodynamic, manufacturing cost, and robustness criteria; performing, in an iterative manner, a sequence of changes to the initial design, by adjusting the initial respective line for the straight line cut for the respective surface of each airfoil to generate different iterative designs; evaluating respective scores for each of the different iterative designs; selecting a design from the initial design and the different iterative designs that generates an optimized score based on the mechanical, aerodynamic, manufacturing cost, and robustness criteria; and cutting along the straight line for the surface of each airfoil, based on the selected design, to form each airfoil.