The gas turbine engine rotor assembly includes a rotor disc with a plurality of sealing tabs projecting radially out from a peripheral surface of the rotor disc in the rear end portion thereof. A said sealing tab is disposed at a tip portion of a fixing member of the disc, formed between pairs of blade root slots, adjacent a trailing edge of the rotor disc. The sealing tabs help to reduce the leakage of secondary air out the back of a blade pocket defined between adjacent blades mounted on the rotor disc.

The present invention is directed to a method for coating a component, wherein the component has a first and a second surface, and wherein the first and the second surface adjoin each other at an edge, in which method i) first of all, the edge between the first and the second surface is rounded, and ii) subsequently, a coating is applied to the first surface.

A method of balancing a gas turbine engine rotor comprises the step of obtaining a rotor disc with a circumferential array of balance tabs projecting from a peripheral rim of the disc. Stress shielding scallops are defined in the rotor disc between the tabs. The balancing is achieved by removing material from at least one of the tabs.

A method of designing a turbine blade includes the steps of forming at least two notches on a tip of a turbine blade, each of the at least two notches having a known dimension. The turbine blade has a pressure side and a suction side. The method further includes the step of operating a gas turbine engine including the turbine blade to expand a length of the turbine blade such that the tip of the turbine engages a casing. The method further includes the steps of viewing the tip of the turbine blade after the step of operating of the gas turbine engine, determining an appearance of the notches on the tip and determining a manufacturing length of the turbine blade based on the step of determining the appearance the notches.

A sheet metal blank is formed into a helical turbine blade by stretching an outer portion of the blank by creating indents and detents thereon. Then the stretched side is flattened and the stretched structure is formed into a helical shape.

The invention concerns a blade (1) designed to be fitted to the rotor of a rotating machine, said blade being characterized in that it exhibits a modular construction including: a rigid, median module (4) formed by a coffer (12), which is itself formed by joining two side panels (13, 14) that are connected to one another by at least one side rail (15), the side panels (13, 14) being connected by at least one weld executed on an internal protuberance on the said side panels that is set back from the intrados (11) and the extrados (1 E) of the blade (1), an upstream module (7) forming a leading edge and added onto the upstream portion of the median module, a downstream module (8) forming a trailing edge and added onto the downstream portion of the median module. Blades for rotating machines and corresponding fabrication methods.

An apparatus for inserting flattened tubes into heat exchanger fins includes a fin stack arranging unit in which heat exchanger fins are stacked in the thickness direction and a guide is inserted into at least a cutaway portion. A flattened tube arranging unit holds flattened tubes in an intermittent arrangement with respect to the cutaway portions into which the guide has not been inserted. A flattened tube insertion driving unit inserts the flattened tubes into the cutaway portions. A contact plate produces a reactive force when the flattened tubes are inserted into the cutaway portions. A flattened tube intermittently inserted fin stack arranging unit aligns the positions of cutaway portions into which the flattened tubes have not been inserted to the positions of the flattened tubes on the flattened tube arranging unit and is used in place of the fin stack arranging unit.

A fan blade of a gas turbine engine is disclosed which is composed of an airfoil having a leading edge and a trailing edge, and a root with a platform and a blade fixing for engaging a fan hub. The fan blade is composed of a core substrate selected from the group consisting of composites and polymers, and an entirety of the airfoil and the root of the fan blade has a nanocrystalline metal outer layer thereon which forms an outer surface fully enveloping the fan blade. The nanocrystalline metal layer formed of a nanocrystalline metal coating has an average grain size of between 10 nm and 500 nm, and the nanocrystalline metal outer layer forms a structural element of the fan blade.

A method of fabricating at least one metal part for a turbine engine, the method including the steps of casting a metal alloy by centrifuging into a permanent metal mold for making a blank; obtaining a cast blank of elongate shape and of section that is circular or polygonal; and machining the blank to make the part.

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.