A turbine engine blade made of composite material including fiber reinforcement obtained by three dimensionally weaving yarns and densified with a matrix, the blade including an airfoil and a blade root forming a single part. The blade root includes two opposite lateral flanks that are substantially plane and that are clamped between two independent pads made of composite material, which pads are fastened against the lateral flanks of the blade root to form a blade root that is bulb-shaped.

An example turbomachine rotor provides a groove that is annular and is configured to receive a root of an airfoil, the groove having a radial cross-sectional area. A ratio of the radial cross-sectional area of the groove to a radial cross-sectional area of the root received within the annular groove is from 2 to 5.

A rotating blade assembly for a turbine engine having a drive shaft, the rotating blade assembly comprising a disc, at least one blade assembly, and a retainer. The disc being operably coupled to the drive shaft and including a seat having at least a portion of a first through hole. The at least one blade assembly having an upper platform, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform. The retainer assembly securing the disc to the at least one blade assembly and comprising a hollow tubular element, a pin, and a fastener.

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 turbine rotor assembly includes a rotor shaft and rotor blades. The rotor shaft includes two protrusions, two bearing surfaces, two first facing surfaces, and two second facing surfaces positioned on the outer side of the first facing surfaces. The blade root section of each of the rotor blades includes two contact surfaces which are contactable with the two bearing surfaces in the radial direction of the rotor shaft, respectively, two first side surfaces each facing a corresponding one of the two first facing surfaces, two second side surfaces each facing a corresponding one of the two second facing surfaces, and two flange sections which are each positioned adjacent to the outer circumferential surface of a corresponding one of the two protrusions.

A rotor disk assembly for a gas turbine engine includes a rotor disk that defines an axis, the rotor disk comprising a slot within a rim of the disk and a spacer within the slot, the spacer having a clearance feature sized to permit rotation of a rotor blade from an insertion position to an installed position.

In the method for carrying out work, according to the invention, in order to maintain a metal turbine engine motor, such as a drum, in an operational state: an upstream and/or downstream face (343) of a circumferential groove (34) of the rotor is ground, with said face having recessed marks that the grinding is intended to remove or at least blur, and foils (43) are placed in the groove between the blades of the rotor (24) and the rotor, at least in the area where the face has been ground.

The present disclosure provides a fir tree coupling for gas turbine engine parts comprising a load beam having a longitudinal axis, a rounded base, a first side, and a second side, wherein the rounded base has a radius of curvature from the first side to the second side, a tooth running parallel to the longitudinal axis and disposed on the first side of the load beam. The fir tree coupling may comprise a channel through the rounded base across a portion of the radius of curvature from the first side to the second side. The channel may comprise a sidewall having a sidewall step cut into a portion of the channel sidewall.

Systems and methods for cooling a rotor assembly disposed within a cavity of an expander fluidly coupled with a cooling source are provided. The system may include an annular body disposed on a rotor disc of the rotor assembly. The rotor disc may also include a plurality of rotor blades mounted thereto via respective roots. The annular body may define at least one fluid passageway fluidly coupling the roots and the cooling source. The annular ring may be configured to substantially prevent mixing of the flue gas with a coolant provided by the cooling source and flowing through the at least one fluid passageway and contacting at least one root. The system may also include a plurality of seal members, each disposed between respective platforms of adjacent rotor blades and configured to substantially prevent the flue gas flowing though the expander from mixing with the coolant.

An assembly for an aircraft turbine engine including an integral element including a disk and a plurality of blades is provided. Each blade has a connection zone connecting the blade to the disk, the connection zone including a first part arranged at the external flowpath delimitation surface provided on the disk, and at least one second end part arranged at a recess formed in the disk along the axial extension of the external surface, the average fillet radius defined by the second end part being larger than the fillet radius defined by the first part. The assembly further includes a flowpath reconstruction part formed in the aerodynamic continuity of the external surface, so as to cover the recess.