Abstract: A one-block bladed disk includes blades, a leading edges of which has sweep and dihedral angles, both strongly variable and increasingly large upon approaching a connection to a hub to improve mechanical strength of the whole and obtain good distribution of gas flow over the height of the blade.

Abstract: A one-block bladed disk includes blades, a leading edges of which has sweep and dihedral angles, both strongly variable and increasingly large upon approaching a connection to a hub to improve mechanical strength of the whole and obtain good distribution of gas flow over the height of the blade.

Abstract: A method of checking turbomachine blades is presented that may be implemented using a computer and a measuring device. Turbomachine blades compatible with embodiments of the method have a profile including a centerline, suction face, pressure face, leading edge and trailing edge. The method measures geometrical coordinates of many points on a blade section profile, calculates an aerodynamic parameter of the blade section as a function of the measured coordinates, verifies whether the calculated aerodynamic parameter value departs from a valid range of parameters from a reference blade, and validates or rejects the blade depending upon whether the value of the aerodynamic parameter falls within the valid range.

Abstract: The present invention relates to the checking of turbomachine blades having a profile comprising a centerline, a suction face, a pressure face, a leading edge and a trailing edge, consisting in: measuring geometrical coordinates of a plurality of points located on the profile of at least one blade section; calculating at least one aerodynamic parameter of the blade section as a function of the measured coordinates; verifying whether the value of the calculated aerodynamic parameter departs from a validity range defined by a value of the nominal aerodynamic parameter of a reference blade and an associated tolerance; and validating the blade if the value of the aerodynamic parameter falls within the validity range or rejecting the blade if the value of the aerodynamic parameter lies outside the validity range. The present invention also relates to a computer program, which can be loaded directly into the memory of a computer, intended to implement the method according to the invention.

Abstract: In a turbomachine compressor stator comprising a plurality of variable-pitch vanes (36), each vane mounted in rotation, about an axis of pivoting (38), in a wall (34) forming a surface of revolution about a longitudinal central axis of rotation of the turbomachine (32), being inclined by a specified angle (&agr;) with respect to a plane perpendicular to this longitudinal central axis, there is provision for this wall to be machined locally opposite an end part of each vane (42) so as to form a spherical portion or pocket (40) whose center is positioned on the axis of pivoting and whose radius is determined so as to ensure, over the entire range of operation of the vane, a predetermined minimum clearance between this end part of each vane and the opposite machined spherical pocket while limiting the depth of machining of this pocket.

Abstract: In a turbomachine compressor stator comprising a plurality of variable-pitch vanes (36), each vane mounted in rotation, about an axis of pivoting (38), in a wall (34) forming a surface of revolution about a longitudinal central axis of rotation of the turbomachine (32), being inclined by a specified angle (&agr;) with respect to a plane perpendicular to this longitudinal central axis, there is provision for this wall to be machined locally opposite an end part of each vane (42) so as to form a spherical portion or pocket (40) whose center is positioned on the axis of pivoting and whose radius is determined so as to ensure, over the entire range of operation of the vane, a predetermined minimum clearance between this end part of each vane and the opposite machined spherical pocket while limiting the depth of machining of this pocket.

Abstract: A separator-diffuser for a gas turbine engine is disclosed wherein the diffuser includes axial walls which function as a guide vane stage for the oxidizer compressor. By incorporating the last, or downstream, guide vane stage into the diffuser structure, the overall axial length and bulk of the gas turbine engine may be significantly reduced. The separator-diffuser has a wall defining a generally annular oxidizer flow passage extending around the central axis of the gas turbine engine, the passage having an oxidizer inlet located downstream of the last rotor stage of the oxidizer compressor so as to enable the oxidizer to pass from the oxidizer compressor into the oxidizer flow passage. A generally annular separator member is located in the oxidizer flow passage downstream of the oxidizer inlet so as to divide the oxidizer flow passage into at least two separate oxidizer flow paths, each flow path having a separate oxidizer outlet.

Abstract: A gas turbine engine impeller having a plurality of blades attached to a rotor disk and an integral, annular collar defining openings through which the blades extend which acts as a platform extending between the adjacent blades. The annular collar is formed as an integral, annular structure and may be formed from a composite material, or other material which is resistant to high temperatures and high centrifugal forces. The annular collar defines a plurality of openings to accommodate the blade portion of the plurality of blades and to enable the blades to extend radially outwardly from the collar. The openings may be formed in the annular collar when the collar itself is formed, or may be machined therein after the collar has been fabricated.