Abstract: A clearance control device for controlling clearance between rotary blade tips and a stationary bushing of a gas turbine having a casing that is provided with at least two annular ridges, the clearance control device including a circular tuning unit that includes air circulation means for circulating air, said means being made up of at least three ducts; air supply means for supplying air to the air flow ducts; and air discharge means for discharging air on the ridges in order to modify the temperature, the air discharge means for each duct being made up of at least one top row having a number N of perforations disposed facing one of the side faces of the ridges and of at least one bottom row having a number 2N of perforations disposed facing a connection radius that connects the ridges to the casing.

Abstract: The invention relates to an axial compressor stator for a gas turbine, the stator having a rigid, external, annular frame and axially juxtaposed rings configured within the frame and bearing annuli of stationary vanes. The rings are defined by arcuate segments affixed to the frame. The inside walls of the arcuate segments externally define the aerodynamic conduit for the compressed gaseous fluid. The arcuate segments are brazed segments that include a honeycomb component sandwiched between an inner sheetmetal bounding the aerodynamic conduit and an outer sheetmetal. The outer sheetmetal solely connects the arcuate segments to the frame.

Abstract: A method and system for detecting damage to the rotor of an aircraft engine using devices for measuring vibration and speed in order to acquire data relating to the speed of the rotor and also to the amplitude and the phase of rotor vibration during a determined flight. The method includes the following steps: reading the acquired data; calculating a mean vibration vector over a determined rotor speed range on the basis of the acquired data; calculating a vector difference between the mean vibration vector of the determined flight and the mean vibration vector of a reference flight for the rotor speed range; comparing the modulus of the vector difference with a predetermined threshold value; and issuing a warning signal when the modulus of the vector difference exceeds the predetermined threshold value, the steps being performed after the determined flight has been completed.

Abstract: A sealing device for sealing between two static portions of a turbomachine, said device comprising a brush gasket formed by a plurality of resilient bristles which are, at least in part, held in a housing, each bristle having a main portion and a secondary portion interconnected by a bent portion, the secondary portion having a free end ensuring leaktight contact with a surface that is to be sealed of the other static portion. The sealing device further comprises a flexible washer of bent section which has a free end bearing both against the free ends of the secondary portions of the bristles of the brush gasket, and against the surface that is to be sealed so that the free ends of the secondary portions of the bristles of the brush gasket are permanently in contact with the surface that is to be sealed.

Abstract: The invention concerns a throttle control device for an aircraft turbine engine. It comprises a control assembly acting on the native command of the turbine engine (MT1-MT3) as a function of a manual input defined by a pilot control element (1). The pilot control element gives a lever angular position signal (CL, 10JS). The control assembly comprises: an automatic device (4) for converting the lever angular position signal into a transformed angular position signal following a selected command law, and an interface (70) for converting the transformed angular position signal into two sinusoidal signals of the resolver type, thus allowing control by the same device of different turbine machines such as turbine machines which have native command by sinusoidal type signals.

Abstract: The cooling system (30) for an aircraft engine (10) comprises a channel (32) that draws off cold air in the secondary air flow (200), and a heat exchanger (34) located in the channel (32) and in which hot air circulates. The channel (32) comprises: a supply pipe (322) and an evacuation pipe (326) fixed to the nacelle (324), an intermediate box (324) located between the supply pipe (322) and the evacuation pipe (326), fixed to the engine (10), and in which the heat exchanger (34) is placed. Application to cooling of hot parts (22) in an aircraft engine (10).

Abstract: The present invention relates to a bearing support for a rotating engine such as a gas turbine engine, with a rotating shaft supported by at least one bearing and a frame, capable of running at least two different running speeds, characterised in that it comprises at least two mechanical supporting structures with different stiffnesses combined so as to support the bearing at one of these stiffnesses. The invention has the advantage that the use of mechanical structures avoids the presence of any hydraulic equipment, since the stiffness forces thus introduced can directly oppose excitation forces generated by the unbalanced mass of the engine shaft.

Abstract: A gas turbine blade for an airplane engine, comprising in its central portion at least a first central cooling circuit comprising at least first and second cavities extending radially on the concave side of the blade, at least one cavity extending on the convex side of the blade, an air admission opening at a radial end of the first concave side cavity for feeding the first central cooling circuit with cooling air, a first passage putting the other radial end of the first concave side cavity into communication with an adjacent radial end of the convex side cavity, a second passage putting the other radial end of the convex side cavity into communication with an adjacent radial end of the second concave side cavity, and outlet orifices opening out into the second concave side cavity and through the concave face of the blade.

Abstract: A ventilation device for a high pressure turbine rotor in a turbomachine, the turbine comprising upstream and downstream turbine disks fitted with blades, the device comprising a cooling circuit being supplied by a cooling airflow D taken from the back of the combustion chamber. The circuit is such that the airflow passes through orifices formed in an upstream flange of the upstream disk, such that this airflow circulates in the axial direction towards the downstream side between an inner reaming of the upstream disk and a downstream flange of the downstream disk, the device also comprising a labyrinth inserted between the two disks, such that the airflow is divided into a first flow F1 and a second flow circulating on each side of labyrinth towards the blades.

Abstract: The invention relates to a turbomachine of axis X including a compressor and an integrated starter/generator. The field magnetic circuit of the starter/generator is mounted in the bore of at least one disk of the compressor and surrounds the secondary magnetic circuit which is secured to a cylindrical shroud provided on the support structure for the bearing of the compressor. The invention applies in particular to single-shaft or multi-shaft turbomachines.

Abstract: A nozzle comprising an inner barrel, a centerbody disposed inside the inner barrel so as to define a first channel for a primary flow, and an outer barrel surrounding the inner barrel so as to define a second annular channel for a secondary flow, the free end of the inner barrel presenting a surface having a first series of undulations made up of negative undulations alternating with positive undulations, the centerbody presenting a second series of undulations made up of negative undulations alternating with positive undulations, the negative and positive undulations of the series of undulations being of radial height that varies angularly and being disposed in such a manner that the negative undulations of the first series are disposed facing negative undulations of the second series, and vice versa, so as to generate turbulence and radial shear between the primary and secondary flows while conserving a constant flow section around the entire circumference of the nozzle.

Abstract: A fan rotor includes a disk having a rim with a plurality of substantially axial grooves that are spaced apart angularly. Removable blades extend radially outwards from the periphery of the disk. Each blade has a blade root in a respective groove. A downstream flange plate can be secured to the disk with the downstream faces of the blade roots being in abutment against the downstream flange plate. An upstream flange plate is secured to the disk so as to retain the blade roots in the grooves. In the rotor, the upstream flange plate is fitted on its downstream face with a resilient device configured to exert sufficient force on the upstream faces of the blade roots, after assembly, to prevent an axial displacement of the blades during normal operation (apart from exceptional events of the type in which foreign bodies are ingested or blades are lost).

Abstract: A gas turbine part comprises a superalloy metal substrate, a bonding underlayer formed on the substrate and comprising an intermetallic compound of aluminum, nickel, and platinum, and a ceramic outer layer anchored on the alumina film formed on the bonding underlayer. The bonding underlayer essentially comprises an Ni—Pt—Al ternary system constituted by an aluminum-enriched ?-NiPt type structure, in particular an Ni—Pt—Al ternary system having a composition NizPtyAlx in which z, y, and x are such that 0.05?z?0.40, 0.30?y?0.60, and 0.15?x?0.40.

Abstract: The present invention pertains to an automated polishing process for semi-finished mechanical parts in titanium or titanium alloy, using a machine with abrasive belt mounted on a tangential contact wheel driven in rotation at a determined speed and applied at a determined pressure, the wheel travelling with respect to the part's surface at a determined rate, characterized by the fact that the abrasive belt consists of superabrasive grains in industrial diamond or cubic boron nitride. The process is applied to geometric conforming of jet engine fan or compressor blades.

Abstract: A ceramic drill bit comprising a cylindrical shank and a shaped portion extending axially in line with the shank and having a free end forming the tip of the bit, a relief face extending from each main cutting edge with a relief angle lying in the range 4° to 10° approximately relative to a plane perpendicular to the axis of rotation of the bit, and two notches forming two secondary cutting faces extending from the central edge with positive cutting angles lying in the range 10 to 70 approximately relative to the axis of rotation of the bit.

Abstract: An axial compressor for a turbomachine is fitted with a device for centripetally bleeding turbine-cooling air. The compressor includes at least two rings of blades, an outer shroud having holes, and a fixed ring of stator vanes placed in the stream between the moving rings of blades. The holes are inlets for the bleed device, opening out into an annular groove beneath the interstice separating the inner platforms of the stator vanes from the rim of the upstream disk. The groove is fitted with fixed air guide devices to impart a centripetal swirling motion to the air flowing therein in the same direction as the compressor so as to reduce the velocity of the air relative to the rotating holes.

Abstract: The pitch control mechanism (controlling the angular position) for stator blades in a turbojet may be adjusted separately for two adjacent stages of blades controlled by a common actuator means 10. To achieve this, the synchronization bar 9 located between the mechanisms leading to the two stages is inclined, this inclination being determined by a pin 22 and groove 23 connection with the casing, while another pin 26 and groove 27 connection is arranged between the synchronization bar 9 and one of the mechanisms. Strongly non-linear displacement laws between the two mechanisms may thus be controlled depending on the shapes and directions of the grooves.

Abstract: The arrangement of bearing supports (20, 50; 22, 52, 520) for a rotating shaft (10) mounted on bearings (20, 50) connected to an engine stator structure (16) by bearing supports 22, 52, 520) of which at least one is made up of N>1 elements (24, 240; 40, 400) effectively acting in parallel to simultaneously connect the said bearing (20, 50) to the said structure (16), of which N?1 of the said. N elements form this connection by means of fusible links (80, 800, 810), is made up of a first bearing (20) supported by a first bearing support (22) and a second bearing (50) supported by a second bearing support (52, 520), that of the two bearing supports (22, 52, 520) which makes up the said N>1 elements being the second bearing support (52, 520). Application to the rotating shaft (10) of an aircraft engine.

Abstract: Blades (3) extending radially from the inner shroud (1) to the outer shroud (2) are attached to the inner shroud by means of a sealing cement containing an organic polyimide-based polymer adhesive.

Abstract: The blade of gas turbine engine compressor of the invention includes a hammer type root to be inserted into a circumferential groove of the rotor of the compressor, a platform (12) integral with the root (11) and supporting an airfoil portion (13), the platform including two edges perpendicular to the axis of the rotor (20, 21) and two curved flanks (22, 23), the curve of the flanks being made out of at least one curve defined by an equation, the curvature centre of the point in the curve whereof the curvature radius is the smallest, being situated within a band (B) central to the platform and accounting for 60% of the width (D) of the platform (12) measured between its parallel rectilinear edges (20, 21), the equation defining the curve being, in the sense of mathematic functions, continuous and with continuous first derivative.