A method is provided for machining an airfoil section (12) of a turbine blade or vane produced by a casting process. The airfoil section (12) has an outer wall (18) delimiting an airfoil interior having one or more internal cooling passages (28). The method involves: receiving design data pertaining to the airfoil section (12), including a nominal outer airfoil form (40.sub.N) and nominal wall thickness (T.sub.N) data; generating a machining path by determining a target outer airfoil form (40.sub.T), the target outer airfoil form (40.sub.T) being generated by adapting the nominal outer airfoil form (40.sub.N) such that a nominal wall thickness (T.sub.N) is maintained at all points on the outer wall around the one or more internal cooling passages (28) in a subsequently machined airfoil section; and machining an outer surface (18a) of the airfoil section (12) produced by the casting process according to the generated machining path, to remove excess material to conform to the generated target outer airfoil form (40.sub.T).

A turbomachine includes a housing assembly having a first housing member with a shroud surface, a bearing housing, and a second housing member. The turbomachine further includes a bearing that supports rotation of a rotating group within the housing assembly. The first housing member has a first axial surface and the bearing housing has a second axial surface that is substantially flush with the first axial surface. The second housing member has a third axial surface facing in an axial direction opposite that of the first and second axial surfaces. The first housing member has a first radial surface and the bearing housing has a second radial surface. The first housing member and the bearing housing are attached to the second housing member. The first and second axial surfaces abut against the third axial surface, and the first radial surface abutting against the second radial surface.

An electric coolant pump for providing an automotive cooling circuit with a coolant includes an electric motor to drive the electric coolant pump, a pump housing, a separating can, a printed circuit board, and permanently height-adjusted supporting elements. The separating can includes a substantially plane separating can bottom wall which lies in a cross plane, and a substantially cylindrical separating can shell which fluidically separates a wet zone from a dry zone. The printed circuit board is arranged substantially parallel to and not in direct contact with the separating can bottom wall so that an axial gap exists which is filled with a heat conductive means. The supporting elements axially support the plane printed circuit board. Each supporting element has a distal tip which is trimmed to define a height-constant axial gap having a nominal gap height between the printed circuit board and the separating can bottom wall.

The invention relates to a method for producing a rotor of a flow engine, namely an integrally bladed rotor with an integral outer shroud, comprising at least the following steps: a rotor blank comprising the integral rotor blades and the integral outer shroud is first produced by means of a generative production method; the rotor blank is then subjected to a separating surface treatment at flow-guiding sections and is subjected, separately therefrom, to a machining surface treatment at non-flow-guiding sections.

A process of forming an aerofoil is provided. The process includes: providing a layered, planar pre-form; inflating and hot creep forming the pre-form to form an intermediate structure having aerofoil pressure and suction surfaces, and a front, edge-receiving portion joining front edges of the pressure and suction surfaces; providing a leading edge piece; and bonding the leading edge piece to the front, edge-receiving portion of the intermediate structure to form an aerofoil in which the leading edge piece forms an aerofoil leading edge.

A method is provided for machining an airfoil section (12) of a turbine blade or vane produced by a casting process. The airfoil section (12) has an outer wall (18) delimiting an airfoil interior having one or more internal cooling passages (28). The method involves: receiving design data pertaining to the airfoil section (12), including a nominal outer airfoil form (40.sub.N) and nominal wall thickness (T.sub.N) data; generating a machining path by determining a target outer airfoil form (40.sub.T), the target outer airfoil form (40.sub.T) being generated by adapting the nominal outer airfoil form (40.sub.N) such that a nominal wall thickness (T.sub.N) is maintained at all points on the outer wall around the one or more internal cooling passages (28) in a subsequently machined airfoil section; and machining an outer surface (18a) of the airfoil section (12) produced by the casting process according to the generated machining path, to remove excess material to conform to the generated target outer airfoil form (40.sub.T).

The present invention relates to a method for machine finishing the shape of a blank casting for a multi-vane, in particular bi-vane, nozzle of a turbine engine, comprising a first vane and a second vane extending substantially in a radial direction between two walls that are radially inner and radially outer, respectively, the suction face of the first vane defining, together with the pressure face of the trailing edge of the second vane, a cross section of flow (SP), the method comprising measuring, by means of probing, the position of predefined points on said respectively radially inner and radially outer walls on the surface of the vanes and calculating the machining allowances (1 and 2 respectively) on the first and second vanes with respect to the theoretical profile at said points, wherein the method comprises calculating said cross section of flow (SP) from the height of the duct between said radially inner and radially outer walls, and values of the machining allowances (1 and 2), a correction of the machining allowance (2) on one of the vanes being applied when the calculated value of the cross section of flow (SP) is outside predefined tolerances.

A plurality of static stator vane are circumferentially spaced in a row axially intermediate the rows of turbine blades. Each stator vane has at least an outer platform with mount structure. Each of the stator vanes are formed of composite materials. The mount structure is provided with sacrificial material. At least a first of the stator vanes is circumferentially adjacent to a second of the stator vanes. An orientation of the first of the stator vanes is re-staggered relative to the second of the stator vanes. The sacrificial material of the first of the stator vanes is machined away such that a final trailing edge of the first of the stator vanes mount structure is now better axially aligned with the trailing edge of the second of the stator vanes mount structure trailing edge.

A method of correcting a manufactured article includes applying a machinable coating to a manufactured ceramic matrix composite article with a manufacturing artifact and machining the machinable coating to provide a desired geometry. A method of correcting a manufactured article and a corrected article are also disclosed.

A verification method for verifying whether the aerodynamic profile of a real blade for an aircraft turbine engine complies with a theoretical blade, the method including constructing a camber line of the theoretical blade and constructing a camber line of the real blade; constructing a relationship for the thickness of the theoretical blade and constructing a relationship for the thickness of the real blade, the thickness relationship of a blade corresponding to the curve plotting the thickness of the blade as a function of curvilinear length along the camber line from a leading edge of the blade to a trailing edge of the blade, where thickness is the dimension of the blade extending perpendicularly to the camber line at each point of the camber line; superposing the thickness relationship of the real blade on the thickness relationship of the theoretical blade; and extracting the leading-edge and trailing edge thicknesses.