A method for manufacturing a propeller for a propeller pump includes providing a base propeller including a hub extending from an axial end in the axial direction, and a plurality of blades fixedly connected to the hub, each blade including a pressure side, a suction side, a leading edge, an initial trailing edge, and a blade tip extending from the leading edge to the initial trailing edge at the end of the blade facing away from the hub, trimming each of the blades of the base propeller the axial direction, and forming a modified trailing edge bye removing a part of the initial trailing edge along the entire pressure side from the hub to the blade tip.

Disclosed is a method of reducing play in a vane arc segment. The vane arc segment includes an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and that extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a radial flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the radial flange, and a thermal insulation element is located adjacent the radial flange. The method includes performing a light scan of the radial flange to produce a digital three-dimensional model of the radial flange, and then machining the thermal insulation element in accordance with the digital three-dimensional model to provide a low-tolerance fit between the radial flange and the thermal insulation element that limits play between the airfoil piece and the thermal insulation element.

A method of forming a component for a gas turbine engine using ceramic matrix composites (CMCs) is provided. The method includes preforming the aerodynamic component into an initial desired shape using the CMCs, executing partial densification of the CMCs, repeating the preforming operations and the executing of the partial densification until a final desired shape of the aerodynamic component is achieved, machining or cutting the CMCs during one or more of the preforming operations and the executing of the partial densification to remove defects from the CMCs and executing a full densification of the CMCs.

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 (Δ1 and Δ2 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 (Δ1 and Δ2), a correction of the machining allowance (Δ2) on one of the vanes being applied when the calculated value of the cross section of flow (SP) is outside predefined tolerances.

Disclosed is a repair method for guide blades of a gas turbine. The method comprises: providing at least one guide blade to be maintained; capturing the actual geometry of the guide blade to be maintained with application of at least one measuring method; comparing the actual geometry captured by the contactless measuring method to a predetermined desired geometry for a corresponding guide blade type; calculating a target geometry for the guide blade to be maintained, which corresponds as much as possible to the desired geometry, such that using optimization parameters, the desired geometry of the guide blade to be maintained is approximated at least in sections along its flow contour; applying material and removing material by machine on the guide blade, such that the calculated target geometry is produced.

An automated technique for finishing gas turbine engine blades or vanes by generating a bespoke tooling path for each blade or vane. The bespoke tooling path is generated by scanning the aerofoil surface to generate a 3-D electronic representation of the surface. The 3-D electronic surface is then analyzed to identify imperfections or defects, and then a machining path a generated through which the imperfections can be removed. The machining path is determined so as to smoothly blend the surface back to the underlying surface where the imperfections had been present. In this way, the resulting aerofoil, once machined, has optimized aerodynamic performance.

Provided is a vacuum pump in which no finish processing has to be carried out after shaping of a cylindrical rotor even in use of a cylindrical rotor obtained by shaping a fiber-reinforced plastic material into a cylindrical shape. The vacuum pump has a turbo-molecular pump section and a thread groove pump section. The upper end section of a cylindrical rotor, which is obtained by shaping a fiber-reinforced plastic material into a cylindrical shape, of the thread groove pump section, is joined to the lower end section of a rotor of the turbo-molecular pump section. A joining portion of the rotor of the turbo-molecular pump section and the cylindrical rotor of the thread groove pump section is disposed upstream of an exhaust passage. As a result, finish processing does not have to be carried out after shaping of the cylindrical rotor. If finish processing is performed after shaping of the cylindrical rotor a resin may be coated onto a rugged portion of the cylindrical rotor, or fibers may be helically wound at a winding angle not greater than 45 degrees.

A vane arc segment includes an airfoil piece that defines first and second platforms and an airfoil section that extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a first platform radial flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the first platform radial flange. A thermal insulation element is situated adjacent the first platform radial flange. The support hardware supports the airfoil piece through the thermal insulation element.

A vane arc segment includes an airfoil piece that defines first and second platforms and an airfoil section that extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a first platform radial flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the first platform radial flange. A thermal insulation element is situated adjacent the first platform radial flange. The support hardware supports the airfoil piece through the thermal insulation element.

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.