Abstract: A turbomachine compressor blade of main radial orientation with respect to a main axis of the turbomachine. The blade includes a radially internal root portion, a radially external tip portion, a radially intermediate portion, a curved portion tangentially in a direction, and at least one straight portion on the root portion and/or on the tip portion.

Abstract: A turbine engine rotor blade including an airfoil defined by plane airfoil sections stacked in a radial direction, each airfoil section being positioned radially at a height H, where height H is expressed as a percentage of the total height of the airfoil, and being identified by its sweep angle ? and its dihedral angle ?, wherein the sweep angle ? varies as a function of height H in such a manner that the sweep angle ? reaches a maximum at a height H?M, the value H?M lying in the range 5% to 40%, and the sweep angle ? increases from 0% to H?M, and wherein the dihedral angle ? varies as a function of H, the dihedral angle ? being a decreasing function of height H for heights H lying in the range 0% to H?1, H?1 lying in the range 10% to 40%.

Abstract: A turbomachine rotor blade including a blade root and a blade tip spaced apart by a blade height, including at least one intermediate segment between the blade root and the blade tip that presents forward sweep over at least 50% of the blade height, and a distal segment with backward sweep between the intermediate segment and the blade tip, the distal segment also presenting a positive tangential angle of inclination.

Abstract: The present invention relates to a method for modelling a part (1), the method being characterized in that it comprises implementing, using data processing means (11) of a piece of equipment (10), steps of: (a) parameterizing a curve of class C1 representing the value of a physical quantity characterizing said part (1) as a function of a position along at least one portion of the part (1), the curve being defined by: a. two end points (PCU0, PCUK) defining the extent of said portion of the part (1); b. at least one intermediate point PCUi, i?[[1,K?1]]) located between the two end points (PCU0, PCUK); c. at least two Bezier curves connected to said intermediate point; the parameterization being carried out using one or more parameters defining said intermediate point; (b) determining optimized values for said parameters of said curve; and (c) returning the determined values to an interface (13) of said piece of equipment (10).

Abstract: The invention relates to an assembly comprising a turbine engine casing and a bladed rotor wheel arranged therein. The casing presents an inside wall having a circumferential strip of abradable material. Facing the tips of the blades, the casing presents upstream the strip of abradable material, and downstream a circumferential groove. The strip of abradable material is defined downstream by the circumferential groove. The downstream end of the circumferential groove is in register with or downstream from the trailing edges of the blades. This arrangement optimizes the use of abradable material in the casings of the turbine engines.

Abstract: A compressor rectifier of a turbomachine including a plurality of stationary blades extending in a circular fashion between an inner shroud and an outer shroud that are concentric and define interblade channels forming an air duct in which air to be compressed flows, the inner shroud including at least one vortex generator extending into the air duct to reduce corner vortices. The vortex generator is positioned axially in the interblade channel, between the axial position of a leading edge of the blades and those of a trailing edge thereof.

Abstract: A turbine engine has at least two successive annular rows of stationary vanes, e.g. formed by the vanes of a nozzle and by an annular row of casing arms arranged downstream from the nozzle Each casing arm extends substantially in a radial plane passing between the trailing edges of two adjacent stationary vanes of the nozzle, and the pitch between these two stationary vanes is greater than the pitch between the other stationary vanes of the nozzle, in such a manner that the wakes formed at the trailing edges of these two stationary vanes pass respectively on either side of the corresponding casing arm.

Abstract: A turbine engine blade comprising an airfoil extending axially between a leading edge and a trailing edge and extending radially between a root and a tip. The leading edge of the airfoil presents a sweep angle that is positive and that increases continuously from the root to a first radial height of the airfoil situated in the range 20% to 40% of the total radial height of the airfoil as measured from the root to the tip, and decreases continuously from this first radial height of the airfoil to the tip.

Abstract: A turbine engine compressor including a casing having its inside wall defining an aerodynamic reference surface for a gas-passing passage and in which a rotor wheel having radial blades is mounted. A circumferential trench is formed in the inside wall of the casing. Its shape is defined from upstream to downstream by three surfaces, respectively an upstream surface, a middle surface, and a downstream surface, which surfaces are substantially conical. The upstream surface extends upstream from the leading edges of the blades. The middle surface is substantially parallel to the aerodynamic reference surface. The downstream surface extends downstream at least as far as the trailing edges of the blades. The junction between the middle and downstream surfaces is in a range of 30% to 80% or 50% to 65% of the axial length of the blades starting from their leading edges.

Abstract: A turbomachine including at least one bladed disk and vortex generators positioned upstream of the blading of the disk is provided. The vortex generators are of variable pitch.

Abstract: A method for designing a multi stage compressor of a turbomachine includes: determining the appropriate number of blades of each rotor blading; determining by instationary computations the trajectories of the slipstreams of the trailing edges of the blades of a rotor blading of an upstream stage n to the leading edges of the blades of a rotor blading of a downstream stage n+1; positioning angularly this rotor blading of the downstream stage n+1 so that the slipstreams pass substantially in the middle of the inter blade circumferential spaces of this blading; repeating these operations for all the stages in order to achieve an aerodynamic coupling on all of the rotor bladings of the compressor; and validating the respective positions of the rotor bladings by new instationary computations on the whole of the compressor.

Abstract: An assembly including an airfoil for a bladed wheel together with a platform, the airfoils in association with such platforms forming a bladed wheel. The platform surface presents a circumferential depression between a leading edge of an airfoil at 60% of the airfoil going downstream. A skeleton curve designates a curve plotting variations in a skeleton angle of the airfoil as a function of position along the axis of the wheel; and a linearized skeleton curve designates a curve that provides a straight line connection between points representing the skeleton angle respectively at 10% and at 90% of an axial extent of the airfoil from its leading edge, and, in a vicinity of the platform, a lowered portion of the skeleton curve lying under the linearized skeleton curve extends axially over at least half of an axial extent of the depression.

Abstract: An assembly including an airfoil and a platform on which the airfoil may be mounted, the surface of the platform presenting a circumferential depression between a leading edge and trailing edge of the airfoil, a deepest section of the depression being situated in an upstream half of the airfoil. A skeleton curve designates a curve of variations of a skeleton angle of the airfoil as a function of position along the axis of the wheel, and a linearized skeleton curve designates a curve of variations of an angle as a function of position along the axis of the wheel, the curve being a straight line connecting together points representing the skeleton angle respectively at 10% and at 90% of the axial extent of the airfoil from the leading edge. In a vicinity of the platform, the skeleton curve presents a raised portion that lies above the linearized skeleton curve.

Abstract: A turbine engine rotor blade including an airfoil defined by plane airfoil sections stacked in a radial direction, each airfoil section being positioned radially at a height H, where height H is expressed as a percentage of the total height of the airfoil, and being identified by its sweep angle ? and its dihedral angle ?, wherein the sweep angle ? varies as a function of height H in such a manner that the sweep angle ? reaches a maximum at a height H?M, the value H?M lying in the range 5% to 40%, and the sweep angle ? increases from 0% to H?M, and wherein the dihedral angle ? varies as a function of H, the dihedral angle ? being a decreasing function of height H for heights H lying in the range 0% to H?1, H?1 lying in the range 10% to 40%.

Abstract: A blade assembly for a turbomachine compressor includes a plurality of individual devices acting on the flow. The individual devices are provided upstream of the blade assembly and are formed at least so as to generate vortices. Each of the individual devices is arranged on an upstream face of a shroud around which a recirculating flow passes, circulating in a cavity. The recirculating flow is reinjected into the principal flow such that the individual devices act simultaneously on the principal flow and on the recirculating flow.

Abstract: The present invention relates to a method for modelling a part (1), the method being characterised in that it comprises implementing, using data processing means (11) of a piece of equipment (10), steps of: (a) parameterising a curve of class C1 representing the value of a physical quantity characterising said part (1) as a function of a position along at least one portion of the part (1), the curve being defined by: a. two end points (PCU0, PCUK) defining the extent of said portion of the part (1); b. at least one intermediate point PCUi, i?[1,K?1]) located between the two end points (PCU0, PCUK); c. at least two Bezier curves connected to said intermediate point; the parameterization being carried out using one or more parameters defining said intermediate point; (b) determining optimised values for said parameters of said curve; and (c) returning the determined values to an interface (13) of said piece of equipment (10).

Abstract: The invention relates to an assembly comprising a turbine engine casing and a bladed rotor wheel arranged therein. The casing presents an inside wall having a circumferential strip of abradable material. Facing the tips of the blades, the casing presents upstream the strip of abradable material, and downstream a circumferential groove. The strip of abradable material is defined downstream by the circumferential groove. The downstream end of the circumferential groove is in register with or downstream from the trailing edges of the blades. This arrangement optimizes the use of abradable material in the casings of the turbine engines.

Abstract: A turbomachine compressor including variable-pitch vanes including an aerofoil section connected by a mounting plate of circular outline to a pivot guided in rotation in an orifice in a casing is disclosed. The mounting plate of the vane includes at least one notch for bleeding air from the compressor stream. The notch is intended to communicate with a hole in the casing in order to remove the air bled off when the vanes are in a first position, and to be closed off by this casing when the vanes are in a second position, so that the flow rate of bled air depends on the pitch angle of the vanes.

Abstract: A compressor rectifier of a turbomachine including a plurality of stationary blades extending in a circular fashion between an inner shroud and an outer shroud that are concentric and define interblade channels forming an air duct in which air to be compressed flows, the inner shroud including at least one vortex generator extending into the air duct to reduce corner vortices. The vortex generator is positioned axially in the interblade channel, between the axial position of a leading edge of the blades and those of a trailing edge thereof.

Abstract: A turbine engine blade comprising an airfoil extending axially between a leading edge and a trailing edge and extending radially between a root and a tip. The leading edge of the airfoil presents a sweep angle that is positive and that increases continuously from the root to a first radial height of the airfoil situated in the range 20% to 40% of the total radial height of the airfoil as measured from the root to the tip, and decreases continuously from this first radial height of the airfoil to the tip.