The invention relates to an aircraft turbine engine (1) comprising a fan (15) as well as a reduction gear (20) including a plurality of rotary elements (52, 58) and driving the fan, the turbine engine also including a gearbox (32) as well as a housing for drawing mechanical power (36) driving the gearbox, the turbine engine comprising a first gear (38) as well as a second gear (40) which is part of the housing for drawing mechanical power (36) and mates with the first gear. According to the invention, the first gear (38) is rotatably secured to a hub (60) of the fan (15).
The invention relates to a system (10) for supplying a turbomachine with fluid, the supply system (10) comprising a low-pressure pumping unit (101) intended to increase the pressure of the fluid flowing toward a downstream circuit (50, 60). According to the invention, the downstream circuit (50, 60) divides at an inlet node (E), situated between the low-pressure pumping unit (101) and the high-pressure volumetric pump (102), into a circuit (60) supplying an injection system (62) and a variable geometries supply circuit (50). The circuit (60) supplying the injection system comprises a high-pressure volumetric pump (102). The variable geometries supply circuit (50) is configured to convey the fluid toward variable geometry devices (54) from the inlet node (E) to an outlet node (S) connecting the variable geometries supply circuit (50) to the upstream circuit (100) between two pumps (101a, 111a) of the low-pressure pumping unit.
The invention relates to an assembly (100) for an aircraft turbine engine, comprising: a rolling bearing bracket (70) defining an inner space (78) on either side thereof; a rotary assembly (58, 60, 38) comprising a first gear (38); a housing for drawing mechanical power (36) comprising a second gear (40) mating with the first gear; a shaft for drawing mechanical power (42), inserted into the housing (36) and rotated by the second gear (40), the shaft (42) passing through a first opening (96) of the bracket (70). According to the invention, the assembly includes means (91) for mounting the housing (36) on the bearing bracket (70), said means passing through a second opening (95) via the bracket (70), said second opening (95) being configured so as to allow the housing (36) to be inserted into the space (78).
The internal cooling of the moving blades of the turbines in aircraft turbomachines is limited in effectiveness because of inhomogeneities of this cooling on each of the pressure-face and suction-face walls. To address this problem, there is proposed a blade comprising a circuit (50) for cooling the airfoil part (34) thereof, in which circuit the cavities interconnected in series are such that the stream of air flows radially toward the outside along the pressure-face wall (40) in pressure-face cavities (52, 56), and radially toward the inside along the suction-face wall (42) in a suction-face cavity (54) that is separated from the pressure-face cavities by an internal wall (58) of the airfoil part. In this way, the force of the Coriolis effect deflects the stream of air toward each of the pressure-face and suction-face walls thereby limiting the inhomogeneity.
The invention relates to a method for synchronising the engines of an airplane, according to an activation logic having a deactivated state, an armed state and an activated state, in which: the switching of the activation logic from the armed state to the activated state is carried out via a first and then a second successive intermediate state; every instance of the activation logic switching from the intermediate state to the activated state involves the following: taking into consideration, on each engine, the activation state of the synchronisation, and exchanging said data between the engines; the switching of the activation logic of one of the engines to the activated state requires that the safety and activation conditions of the other engines are all met; if one of the engines enters the deactivated state, the others do so as well; and, for each engine, the switching of the activation logic from the armed state to the first and then to the second intermediate state takes place automatically when a first portion and then a second portion of the safety and/or activation conditions are met.
A fuel gear pump (4') works to supply a pre-defined flow rate, but with low or zero pressure rise. In order to ensure the hydrodynamic lifting of bearings (19) supporting the pinions (11), or some of them, a sealing lining (46) is added between them, in order to delimit a closed cavity (47), in order to ensure lifting by a fluid having more suitable viscosity properties, instead of the lifting being ensured by the fuel itself which is pumped. Possible application to fuel pumps of aircraft engines, in which the pump (4') is a high-pressure pump combined with a low-pressure pump.
F04C 2/08 - Rotary-piston machines or pumps of intermeshing-engagement type, i.e. with engagement of co-operating members similar to that of toothed gearing
F02C 7/236 - Fuel delivery systems comprising two or more pumps
F04C 11/00 - Combinations of two or more machines or pumps, each being of rotary-piston or oscillating-piston type; Pumping installations
F04C 15/00 - Component parts, details or accessories of machines, pumps or pumping installations, not provided for in groups
7.
BLADE FOR A TURBINE ENGINE PROPELLER, IN PARTICULAR A PROPFAN ENGINE, PROPELLER, AND TURBINE ENGINE COMPRISING SUCH A BLADE
The invention relates to a blade (11A) for a turbine engine propeller, in particular a propfan engine, comprising a protruding part (16) on the leading edge (17) thereof, characterised in that said blade comprises means for controlling the position of the protruding part along the leading edge thereof.
Mandrels (1, 2) for the rolling of a ring (7) comprise impressions (5, 6) to accommodate the ring, which impressions have fillet radii where the various faces meet, so as to avoid the creation of sharp corners on the rolled ring (7) and the accidental formation of cracks.
The invention concerns a turbomachine bearing housing (E) comprising a fixed wall (9), a rotating shaft (5), first and second seals (10, 20) between the wall and the shaft, and a chamber (Cam) between the fixed wall (9) and a stator element (19) supplied with air via an opening (19a) close to said shaft (5). The housing is characterised by the fact that an air guide means (30) is arranged along the surface of the wall (9) of the housing, outside same, such that at least a portion of the air exiting the guide means passes between the first seal (10) and the shaft, said guide means being supplied with air by an air intake separated radially from the shaft, the air from the air intake being at a higher pressure than at the shaft.
F01D 11/04 - Preventing or minimising internal leakage of working fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
The invention concerns a turbomachine (1) characterised in that it comprises: - an exhaust housing (7), comprising a plurality of arms (10), the space separating the arms defining openings (13) in which there circulates a primary air flow (29) of the turbomachine (1), - at least one conduit (2), a) configured to collect a compressed air flow at one of the ends (3) of same, b) the other end of the conduit (2) being connected to at least one opening (13) of the exhaust housing (7), so as to insert the collected air flow into said primary air flow (29), said collected air flow having, when inserted into the opening (13), a Mach number less than or equal to 0.5.
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
11.
MOBILE MEMBER OF A TURBOMACHINE WHICH COMPRISES MEANS FOR CHANGING THE RESONANCE FREQUENCY OF SAME
The invention proposes a rotor (10) of an aircraft turbomachine having a main axis A, which comprises means (14) for modifying the critical speed of the rotor (10) depending on whether the rotational speed of the rotor (10) is lower or higher than a predefined rotational speed, comprising: a component (16) that is capable of occupying a first state or a second state depending on whether the rotational speed of the rotor (10) is lower or higher than the predefined rotational speed, each state of the component (16) corresponding to a critical speed of the rotor (10); and means (18) for driving the component (16) to one or the other of the two states thereof, depending on the rotational speed of the rotor (10), characterised in that the means (14) for modifying the critical speed of the rotor (10) further comprise a component (38) that engages with the drive means (18) and is capable of being deformed elastically between one or the other of two stable forms, each of which corresponds to a state of said component (16).
The invention relates to a method for fire protection (S) of a part (1) of a gas-turbine engine made of a composite material comprising a main fibrous reinforcement compregnated by a main matrix, the protection method (S) comprising the following steps: preforming (S1) a panel of prepreg (20) such as to grant same a shape corresponding to the shape of a surface (3) of the part (1) to be protected against fire, said panel of prepreg (20) comprising a secondary fibrous reinforcement compregnated by a secondary matrix; applying (S2) the panel of prepreg (20) thus preformed to the part (1); and securing (S3) the panel of prepreg (20) to the surface (3) by thermal treatment of the part (1) provided with said panel of prepreg (20) in order to obtain a fire-protection layer (2).
Method (S) for protecting against fire a fan casing (1) comprising a roughly cylindrical barrel (10) having a main direction extending along a longitudinal axis (X) and an upstream flange (20) extending radially with respect to the longitudinal axis (X) from an upstream end of the barrel (10), the fan casing (1) being made of a composite comprising a fibrous reinforcement densified by a matrix, said matrix being polymerized, the protection method(S) comprising the following steps: - laying (SI) widths containing glass fibre pre-impregnated with a resin capable of affording the fan casing with thermal protection against fire on an upstream radial face (22) of the upstream flange (20), and - polymerizing (S2) the resin in order to obtain a protective layer (2).
The invention pertains to a method for monitoring an aircraft engine retractable doors thrust reverser, the thrust reverser being a reverser having hydraulic actuators and being provided with contactors (3a, 4a, 5a, 3B, 4b, 5b, Sa, Sb) arranged so as to each return an item of information about the position of the doors, the engine comprising a computer (3) configured to carry out measurements (E1) of a parameter representative of the position of the contactors on the basis of the information returned by the contactors, characterized in that it comprises a calculation (E2) of one or more statistical indicators of the parameter measured and an analysis (E3) of the temporal evolution of the statistical indicator or indicators calculated. The invention extends also to a computer program for the implementation of this method.
The present invention relates to an epicyclic reduction device (70) for rotating a first set of blades of a turbomachine, comprising a sun gear (74) centred on a longitudinal axis (12) of the turbomachine and connected to a rotor (76) of the turbomachine so as to be rotated; at least one planet gear (78) meshing with the sun gear; a planet gear carrier (80) rotationally bearing the planet gear and connected to a first set of blades (82) to rotate same; and an annulus gear (72) meshing with the planet gear; the sun gear being connected to the rotor by a first ball-type constant velocity joint (84).
F02K 3/072 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with counter-rotating rotors
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
16.
AIRCRAFT PROPULSION ASSEMBLY WITH FIRE EXTINGUISHING SYSTEM
Aircraft propulsion assembly (10) comprising an engine (16), a nacelle (18) surrounding the engine, and a system for extinguishing a fire that may occur in the engine and/or in the nacelle, this extinguishing system comprising means (34) for supplying an extinguishant to at least one extinguishant distribution pipe (36) which opens into a cavity (32) of the engine and/or a cavity (26) of the nacelle, characterized in that it further comprises means (48) for supplying said at least one pipe with air so as to ventilate the or each cavity.
The invention relates to an aircraft turbomachine (10) comprising a nacelle and an engine (12) comprising at least one outflowing jet of air, characterised in that a heat exchanger (20) of the precooler type for supplying air to the aircraft is mounted in the nacelle, said exchanger comprising a primary circuit, the inlet of which is connected to means for taking compressed air from the engine and the outlet of which is connected to means for supplying air to the aircraft, and a secondary circuit supplied with air taken from said air flow.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
Inventor
Schneider, Julien
Hild, Francois
Leclerc, Hugo
Roux, Stephane
Abstract
The invention relates to a method for characterising a part (10), including a step of obtaining an X-ray tomography image of the part, followed by a step (200) of correlating said image with a reference (20), characterised in that the correlation step (200) includes searching, in a predefined set (30) of transformations of X-ray tomography images, for a transformation (40) that minimises the difference (50) between the image and the reference in order to characterise (300, 350) the inside of said part (10).
G01N 23/046 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and forming images of the material using tomography, e.g. computed tomography [CT]
19.
HOUSING MADE FROM AN ORGANIC-MATRIX COMPOSITE MATERIAL PROMOTING THE DISCHARGE OF SMOKE
A gas turbine housing (100) made from an organic-matrix composite material comprising a reinforcement densified by an organic matrix delimits an inner volume. The housing comprises, on the inner face (101) of same, a structural portion (120) having a first face (120a) facing the inner face of the housing and a second opposing face (120b) defining a flow channel portion (102). Recesses (130) opening into the inner volume of the housing are present between the inner face (101) of the housing and the first face (120a) of the structural portion (120) facing said inner face of the housing. The recesses (130) allow the gases produced by the degradation of the resin of the housing in case of a fire to be discharged from the flow channel side.
F02C 7/05 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles
The invention relates to a method for manufacturing a part coated with a protective coating, the method comprising the following step: forming a protective coating on the outer surface of a part by micro-arc oxidation treatment, the part comprising a niobium matrix which contains metal-silicide insertions, the current passing through the part being monitored during the micro-arc oxidation treatment in order to subject the part to a series of current cycles, the ratio (amount of positive charge applied to the part) / (amount of negative charge applied to the part) being, for each current cycle, 0.80 to 1.6.
The invention proposes a guide arm for guiding at least one element having an elongated shape (20), corresponding to a set of cables and/or pipes. The arm comprises an inner cavity (62) opening on the outside of the arm at each of the ends thereof, and in which the elements having an elongated shape can extend. According to the invention, this structure more particularly comprises - a frame (8) comprising a beam (18) linked to means (30, 36, 44) for holding the elements having an elongated shape on the outside and along the beam, and - a cover (54) of which the walls (56, 58) cover the holding means of the frame, and are engaged with the beam, in such a way as to form the inner cavity in which the elements having an elongated shape extend, shock-absorbing means (68) being arranged between the means (30, 36, 44) for holding the elements having an elongated shape (20), and the longitudinal walls (56, 58), in such a way as to reduce and damp the movements of the means for holding in the cavity (62).
F01D 9/06 - Fluid supply conduits to nozzles or the like
F16L 3/10 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets substantially surrounding the pipe, cable or protective tubing divided, i.e. with two members engaging the pipe, cable or protective tubing
F16L 3/233 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets specially adapted for supporting a number of parallel pipes at intervals for a bundle of pipes or a plurality of pipes placed side by side in contact with each other by means of a flexible band
The invention relates to a device (100) for cleaning, and in particular for degritting or desanding, a turbomachine module (110), characterised by comprising: (i) means (102, 104) for isolating bearings of the module, by containment in a closed enclosure (106); (ii) means (112) for overpressurising said enclosure; (iii) means (114) for stripping material deposited in the walls of annular recesses of the module, for example by spraying compressed air onto said walls; and (iv) means (116) for sucking up the material thus stripped.
The invention relates to a turbine engine compressor, in particular of an aeroplane turboprop or turbofan, including a stator comprising an annular casing and at least one annular row of variable-pitch vanes, each vane comprising a radially external end including a pivot mounted in an opening of the casing and connected by a linking member to a control ring (38) capable of pivoting axially relative to the casing, the linking member comprising a first end attached to the pivot of the vane and a second end comprising a pin inserted in a hole (52, 58) of the control ring (38), characterised in that at least one (58) of the holes (52, 58) of the control ring (38), which is used for inserting the pins of the linking members, has an oblong shape and extends in the circumferential direction such as to enable the pin to move into said oblong hole (58), during the rotation of the control ring (38).
The present invention relates to a turbomachine component (1) or collection of components comprising at least a first and a second blade (3I, 3E) and a platform (2) from which the blades (3I, 3E) extend, characterized in that the platform (2), between the pressure face of the first blade (3I) and the suction face of the second blade (3E) has a non-axisymmetric surface (S) defining a plurality of fins (4) of substantially triangular section extending downstream of a leading edge (BA) of each of the blades (3I, 3E), each fin (4) being associated with a leading position and a trailing position on the surface (S), between which positions the fin (4) extends, such that: the leading position is situated at between 5% and 35% length relative to a chord of the blade (3I, 3E) extending from a leading edge (BA) to a trailing edge (BF) of the blade (3I, 3E); - the further a fin (4) is from the suction face of the second blade (3E), the further the leading position of said fin (4) is axially from the leading edge (BA) of the blades (3I, 3E).
Blade intended for a turbomachine impeller comprising N blades. At one end, the blade has a platform formed integrally with an airfoil of the blade. Over part of the axial extent of the blade, a section on a plane perpendicular to the axis (X) of the impeller of the flow path of the platform consists mainly of two straight-line segments arranged respectively on the two sides of the airfoil. These segments form, on each side of the airfoil, an angle of 90°-180°/N with respect to the radial direction. Figure 3.
The present invention relates to a turbomachine component (1) or collection of components comprising at least a first and a second blade (3I, 3E) and a platform (2) from which the blades (3I, 3E) extend, characterized in that the platform (2) has a non-axisymmetric surface (S) bounded by a first and a second end plane (PS, PR) and defined by at least two class C construction curves each one representing the value of a radius of said surface (S) as a function of a position between the pressure face of the first blade (3I) and the suction face of the second blade (3E) in a plane substantially parallel to the end planes (PS, PR), these including at least one upstream curve and one downstream curve; each construction curve being defined by at least one pressure face control end point and one suction face control end point such that: - the tangent to the downstream curve at the suction face control end point 20 is inclined by at most 5°; - any other tangent to a construction curve at a control end point is inclined by at least 5°.
Blade for a turbo machine impeller comprising a root, an air foil and a tip. The root and the tip comprise platforms having surfaces (15) on the side of the air foil, the surfaces respectively being referred to as the root and tip flow path. Each of these flow paths is made up of a pressure-face part and of a suction-face part which are situated respectively on the side of the pressure face and of the suction face and are separated by a crest curve (45, 65). Blade manufacture is made easier notably by virtue of the fact that any point on a first surface out of the pressure face and the suction face and any point on the root and tip flow path parts situated on the side of the first surface has a normal that makes an acute or right angle with respect to a direction referred to as the first direction of manufacture. Method for modelling the blade.
Aircraft propulsion unit comprising a motor and a nacelle comprising a casing (16) of revolution delimiting a flow vein of an air flow, characterized in that this casing comprises at least two openings closed by removable and interchangeable panels (18), at least one of these panels carrying equipment (24) of the propulsion system.
The invention concerns a method and system for forecasting maintenance operations to be applied to an aircraft engine comprising a plurality of elements monitored by damage counters, each damage counter being limited by a corresponding damage ceiling, characterised in that it comprises: - processing means (7) suitable for simulating a consumption of said damage counters (C1-Cm) by iteratively pulling a series of simulation missions from a learning database (9) containing test missions, - processing means (7) suitable for determining, at each iteration, an accumulation of consumption of each of said damage counters until at least one counter counting damage related to a current simulation mission reaches the damage ceiling associated with same, - processing means (7) suitable for applying a maintenance strategy to said current simulation mission to determine maintenance indicators representative of the maintenance operations to be planned on the aircraft engine.
The invention concerns a cutting table (100) for cutting a fibrous preform obtained by three-dimensional weaving and comprising two portions that are linked together by at least one separating area and that have contours of different shapes, the cutting table comprising a plate (104) provided with a cavity (108) intended to receive, flat, one of the portions of the preform to be cut, sacrificial plates (110) intended to be interposed between the portions of the preform to be cut and to be secured to the plate, at least one cutting template (114) intended to be applied to the portion of the fibrous preform that is not positioned in the cavity, and means (118) for applying a compacting pressure to the cutting template. The invention also concerns a method for cutting a fibrous preform using such a cutting table.
A method for impregnation of a fibrous preform (10) by an impregnation composition (20), the method comprising the following step: a) application of a liquid (30) onto a structure, the structure comprising: - a chamber (2) in which a fibrous preform (10) to be impregnated is present, the chamber (2) being defined between a rigid support (3) on which the fibrous preform (10) is placed and a wall (4), the wall (4) comprising a face (4a) located facing the fibrous preform (10), and - an impregnation composition (20), intended to impregnate the fibrous preform (10), the impregnation composition being present in the chamber (2), the liquid (30) being applied on the wall (4) of the opposite side of the chamber (2), the wall (4) being configured so that the face (4a) located facing the fibrous preform (10) retains its shape during application of the liquid (30), the applied liquid (30) enabling creation of a sufficient pressure to displace the wall (4) towards the rigid support (3) and impregnating the fibrous preform (10) with the impregnation composition (20).
B29C 70/44 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
B29C 43/10 - Isostatic pressing, i.e. using non-rigid pressure-exerting members against rigid parts or dies
B30B 1/00 - Presses, using a press ram, characterised by the features of the drive therefor, pressure being transmitted directly, or through simple thrust or tension members only, to the press ram or platen
32.
DRAINED FLUID EVACUATION STUB FOR A PROPULSION ASSEMBLY
Drained fluid evacuation stub (16) for a propulsion assembly (10), comprising a drained fluid storage cavity and at least one orifice (32) for evacuation of the fluids contained in said cavity, characterized in that it comprises means (36, 38) for detecting a pressure difference with the exterior of the stub and a component for purging the cavity which is movable between a first closed position of the evacuation orifice and a second release position of the orifice, the component being configured to move from the first to the second position when the pressure difference is greater than or equal to a predetermined value.
The invention relates to a device (110) for retaining drained fluids for a propulsive assembly, comprising a cavity for storing the drained fluids and two walls (118, 120) mounted at the opening of said cavity, the cavity having a fluid storage volume V1 when the device is in a substantially vertical position, and each wall being configured such as to define a fluid storage volume (V2 and V3 respectively) in the cavity when the device is in a substantially horizontal position, each of the volumes V2 and V3 being at least equal to the volume V1. The invention also relates to a propulsive assembly comprising a device for retaining drained fluids.
Motive flow valve for aircraft engine comprising: a shaft comprising high-pressure and low-pressure chambers, a fuel inlet and outlet, a shoulder arranged between the inlet and outlet, defining an abutment surface, a drawer displaceable relatively to the shaft under a pressure difference between the high- and low-pressure chambers, between a closed and open position wherein the drawer obstructs and releases the fuel inlet, the drawer comprising a portion extending into an intermediate chamber, and a channel exiting on the one hand in a first portion of the intermediate chamber and communicating with the fuel outlet in the open and closed positions, and on the other hand in a second portion of the intermediate chamber that only communicates with the fuel inlet or outlet via the channel in the open and closed positions, and a sealing element interposed between the abutment surface and the drawer in the closed position.
F16K 3/30 - Gate valves or sliding valves, i.e. cut-off apparatus with closing members having a sliding movement along the seat for opening and closing - Details
B64D 37/00 - Arrangements in connection with fuel supply for power plant
F02C 7/232 - Fuel valves; Draining valves or systems
F16K 3/26 - Gate valves or sliding valves, i.e. cut-off apparatus with closing members having a sliding movement along the seat for opening and closing with sealing faces shaped as surfaces of solids of revolution with cylindrical valve members with fluid passages in the valve member
F16K 17/04 - Safety valves; Equalising valves closing on insufficient pressure on one side spring-loaded
35.
BALANCED TURBINE ENGINE PORTION AND TURBINE ENGINE
The invention relates to a balanced turbine engine portion. Said portion comprises at least one angular section (21) arranged such as to form a balancing ring (20) centred on a ring axis (C). Said angular section (21) comprises a plurality of attachment elements (30), a bearing surface (5) with a shape that matches the balancing ring (20), the angular section (21) abutting with said bearing surface (5). Said portion also comprises a plurality of balance weights (40), each attached to the corresponding attachment element (30) of the angular section (21), at least one of said balance weights also being useful as an attachment means for attaching the angular section (21) to the bearing surface (5). The invention also relates to a turbine engine comprising such a balanced portion.
The invention relates to a device (10) for centring and guiding the rotation of a turbine engine shaft, in which the outer ring (18) of a bearing is retained axially upstream and downstream by retaining means (52, 72) that engage with a bearing mounting (20) and with coupling means (29) including resiliently deformable means (32) connecting the outer ring to the bearing mounting, said retaining means being separate from a binding band (28) of the device. The invention also provides a method for assembling such a device in which the retaining means (52, 72) are pre-assembled with the coupling means (29) prior to the final assembly of the coupling means with the bearing mounting (20). The device and the method have the combined advantages of axially retaining the outer ring in two opposing directions and having a particularly straightforward assembly.
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
37.
DEVICE FOR GUIDING SYNCHRONIZING RING VANES WITH VARIABLE PITCH ANGLE OF A TURBINE ENGINE AND METHOD FOR ASSEMBLING SUCH A DEVICE
The invention relates to a device for guiding synchronizing ring vanes with variable pitch angle of a turbine engine, including a plurality of angular inner ring sectors placed end-to-end to form an inner ring (26), each inner ring sector including shafts (24) passing radially from one side of the inner ring sector to the other, a plurality of cylindrical bushes (22) which are each mounted in a shaft of the inner ring from the inside and which are each intended for receiving a guiding pivot (12) of a synchronizing ring vane (4), a plurality of angular reconstitution ring sectors which are placed end-to-end to form a reconstitution ring (36) and which are mounted radially from the inside on the inner ring, and a plurality of locking elements passing axially through the inner and reconstitution rings such as to assemble said rings together. The invention also relates to a method for assembling such a device.
The invention relates to a fan, in particular for a small turbine engine such as a jet engine, having a hub ratio corresponding to the ratio of the diameter of the inner limit of the air intake section (26) at the radially internal ends of the leading edges of the fan blades (10), divided by the diameter of the circle through which the outer ends of the fan blades pass, which has a value of 0.25 to 0.27.
F01D 5/32 - Locking, e.g. by final locking-blades or keys
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream (26) at the radially inner ends of the leading edges of the fan blades (10), divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.
The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream (26) at the radially inner ends of the leading edges of the fan blades (10), divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.
The invention relates to a turbofan engine comprising a fan (S) driven, via a fan shaft (3) supported by at least two first bearings (11, 12), by a turbine shaft (4) supported by at least one second bearing (10) comprising a stationary ring (25) and a movable ring (26), said turbine shaft driving said fan shaft (3) through a device for reducing the speed of rotation (7), said device for reducing the speed of rotation and said first and second bearings being housed in a lubrication enclosure (E1) in which the shell comprises stationary portions and movable portions connected to one another by sealing means (29, 30, 31), said reducing device comprising an inducer (27) shaped so as to receive the torque transmitted by said turbine shaft via driving means (8, 9) connected to said movable ring, characterised in that the lubrication enclosure forms a coaxial ring with the turbine shaft and said driving means (8, 9) comprise a girth gear which is part of the movable sealing walls of the shell of the lubrication enclosure (E1).
F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
F02C 3/067 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages having counter-rotating rotors
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
42.
MODULAR ENGINE, SUCH AS A JET ENGINE, WITH A SPEED REDUCTION GEAR
The present invention relates to an engine (1) with a modular structure comprising a plurality of coaxial modules (A, B, C) with, at one end, a first module (A) comprising a power transmission shaft (3) and a speed reduction gear (7), said power transmission shaft being driven via the speed reduction gear (7) by a turbine shaft (2) secured to one (C) of said coaxial modules that is separate from the first module, the speed reduction gear comprising a drive means (8 and 9) fixed to the turbine shaft (2) and to a journal (13) of a shaft of a low-pressure compressor rotor (1 a), characterized in that it comprises a first nut (16) for fastening the drive means to the journal and a second nut (14) for fastening the drive means to the turbine shaft.
A YSZ-type ceramic layer is deposited on a tie sublayer by thermal spraying using a plasma arc torch, said tie sublayer being itself deposited on the part to be protected. A sintering post treatment is carried out by means of a sweep of the ceramic layer by the beam of the plasma arc torch, the temperature at the point of impact of the beam at the surface of the ceramic layer (C) being, during this sweep, between 1300°C and 1700°C.
C23C 4/12 - Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the method of spraying
The invention relates to a turbomachine (30) comprising: a compressor stage and a turbine stage, each stage comprising at least one disk (42); and a tubular shaft (31) sleeve (33) extending along the axis (32) of the turbomachine, wherein the sleeve (33) comprises at least one tab (40) extending from an outer radial surface (41) of the sleeve and facing the disk (42), the tab (40) being designed to come into contact with the disk (42) when the sleeve (33) is in rotation about the axis (32) of the turbomachine.
The present invention relates to a bypass engine bearing holder (1) that holds an upstream bearing (6) and defines, with said upstream bearing, an oil chamber (100) and an air chamber (200), comprising a frusto-conical portion (11) defining an upstream bearing chamber (160) and a downstream inner chamber (150), and comprises an outer collar (13) connected, by a weld (135), to a flange (15) that extends outward from the frusto-conical portion (11). The outer collar (13) has a sealable gimlet (131) engaging with the upstream bearing (6) such as to seal the oil chamber (100). The bearing holder (1) comprises a plurality of oil recovery ducts (8) leading to the downstream inner chamber (150) and to the upstream bearing chamber (160). The oil recovery ducts (8) lead to the upstream bearing chamber (160), downstream from the weld (135) of the outer collar (13) on the flange (15), the weld (135) of the outer collar (13) being axisymmetric.
The invention relates to a multi-point fuel injection device (1) for an aircraft engine (M), comprising an inlet line (10), at least two injection lines (11, 12), and a purge line (14), a fuel distributor member (2) connected to each line and comprising a moveable element (22) which comprises an injection passage (223), in which the moveable element (22) additionally comprises a purge passage (226), and is configured to adopt a first range of positions in which the injection passage (223) interconnects the inlet line (10) and the injection lines (11, 12), and a second range of positions in which the injection passage (223) interconnects the inlet line (10) and at least a first injection line (11) while the purge passage (226) interconnects the purge line (14) and at least a second injection line (12), the device being characterized in that it additionally comprises an actuator adapted to move the moveable element into a safety position when a failure of the distribution member is detected, the injection passage (223) interconnecting, in this safety position of the moveable element (22), the inlet line (10) and the first injection line (11) while the purge passage (226) does not interconnect the purge line (14) to any of the injection lines (12).
F23R 3/34 - Feeding into different combustion zones
F02C 7/228 - Dividing fuel between various burners
F02C 7/232 - Fuel valves; Draining valves or systems
F02C 9/34 - Joint control of separate flows to main and auxiliary burners
F02M 41/04 - Fuel-injection apparatus with two or more injectors fed from a common pressure-source sequentially by means of a distributor the distributor being spaced from pumping elements the distributor reciprocating
F23D 11/00 - Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space
F23K 5/06 - Liquid fuel from a central source to a plurality of burners
47.
SEALING SYSTEM WITH TWO ROWS OF COMPLEMENTARY SEALING ELEMENTS
The invention concerns a sealing system, in a cavity (C) under a stator (10), of a turbomachine vein (VC, VT), the cavity (C) being located between a stator (10) vane (PS) root (SI) and an additional rotor member (11), the root (SI) comprising two surfaces (21, 24a) each provided with an abradable coating (22, 32), the rotor member (11) being provided with first and second sealing elements (23, 33), disposed respectively facing the first and second surfaces (21, 24a), the first surface (21) and the first sealing element (23) forming a first sealing pair (20) and together delimiting a first leakage section, the second surface (24a) and the second sealing element (33) forming a second sealing pair (30) and together delimiting a second leakage section, one of the two pairs (20, 30) moving to a minimum leakage section when the other (30, 20) moves to a maximum leakage section, and vice versa.
The invention relates to a device (1) for replacing machining inserts (900) on a tool (800) including a body (810) and a head (820) supporting at least one machining insert (900), each insert being maintained on the head (820) by a screw. The device (1) includes: a positioner (50) comprising a supporting element (59) capable of supporting the body (810) of the tool; a screwing station having a screwdriver (60) capable of screwing and unscrewing the screws, the positioner (50) being capable of moving the tool relative to the screwing station; a gripping device (70) capable of gripping and placing an insert (900); a conveyor (500) comprising a plurality of insert containers (510), along which a first station (100) having the gripping device (70), a second station (200) having a mechanism (20) for rotating the inserts on the axes thereof, a third station (300) having a mechanism (30) for unloading the inserts, and a fourth station (400) having a mechanism (40) for supplying inserts are distributed; a transport mechanism (80) capable of moving the gripping device (70) between the positioner (50) and the first station (100) of the conveyor (500); and a control centre (600) capable of automatically controlling one or more of the mechanisms and devices and/or the conveyor of the replacement device (1).
The invention relates to a turbine engine, comprising two structural annular casings (16, 22) connected to one another by means (40, 54) for absorbing stresses from the thrust of the engine, which include connecting rods (54), characterised in that said thrust-absorbing means also include at least one accessory gearbox (40) which is attached to a first one of said casings (16) and which is connected by said connecting rods to the other one of said casings (22).
A fiber preform for a hollow turbine engine vane, the preform comprising a main fiber structure obtained by three-dimensional weaving and including at least one main part (41), wherein the main part (41) extends from a first link strip (44p), includes a first main longitudinal portion (46) suitable for forming essentially a pressure side wall of an airfoil, then includes an U-turn bend portion (45) suitable for forming essentially a leading edge or a trailing edge of the airfoil, then includes a second main longitudinal portion (47) facing the first main longitudinal portion (46) and suitable for forming essentially a suction side wall of the airfoil, and terminating at a second link strip (44q), wherein the first and second link strips (44p, 44q) are secured to each other and form a link portion (44) of the main fiber structure, and wherein the main longitudinal portions (46, 47) are spaced apart so as to form a gap between said main longitudinal portions (46, 47) suitable for forming a hollow in the airfoil.
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
51.
METHOD FOR MONITORING A LOCKING SYSTEM FOR A TURBINE ENGINE THRUST REVERSER
The invention relates to a method for monitoring a locking system (1) comprising N locks (2a, 2b, 2c, 2d, 2e). Each lock (2a, 2b, 2c, 2d, 2e) is monitored by two locking sensors (3a, 4a, 3b, 4b, 3c, 4c, 3d, 4d, 3e, 4e). Each locking sensor is capable of indicating if the lock that said locking sensor is monitoring is in a locked or unlocked state. Each locking sensor can be in a valid or invalid state. The method comprises the following steps: - determining the state of the locking system (1) on the basis of the state of the locks detected by the locking sensors; and - determining a reliability level associated with the state of the locking system on the basis of the number of valid locking sensors monitoring the locks that are in the same state as the locking system.
SOCIETE LORRAINE DE CONSTRUCTION AERONAUTIQUE (France)
Inventor
Poisson, Mathieu Ange
Orcel, Stephane
Glemarec, Guillaume
Pacary, Jean-Luc
Abstract
Turboprop air intake A turboprop (110), comprising a rotary propeller (112) upstream from an engine (114) and an air intake (116) that is not coaxial to the propeller, said air intake defining a conduit (119) for supplying air to the engine and further defining a bypass (124) to said conduit, the bypass comprising an outlet (126) oriented substantially axially towards the downstream of the engine, the turboprop further comprising a nacelle (130) surrounding the engine and the air intake, characterised in that the air intake is secured to a housing (123) of the engine and is not rigidly connected to the nacelle, so as to allow, during operation, relative movements between the air intake and the nacelle, said outlet (126) being connected by a flexible link (140) to an intake of an air circuit carried by the nacelle.
The present invention concerns a turbomachine part (1) comprising at least first and second blades (3, 31, 3E), and a platform (2) from which the blades (3, 31, 3E) extend, characterised in that the platform (2) has a non-axisymmetric surface (S) limited by first and second end planes (PS, PR), and defined by at least three construction curves (PC-A, PC-C, PC- F) of class C1 each representing the value of a radius of said surface (S) on the basis of a position between the lower surface of the first blade (31) and the upper surface of the second blade (3E) according to a plane substantially parallel to the end planes (PS, PR), including: - a first curve (PC-C) that increases in the vicinity of the second blade (3E); - a second curve (PC-F) disposed between the first curve (PC-C) and a trailing edge (BF) of the first and second blades (3, 31, 3E), and that decreases in the vicinity of the second blade (3E); - a third curve (PC-A) disposed between the first curve (PC-C) and a leading edge (BA) of the first and second blades (3, 31, 3E), and having a minimum at the second blade (31).
The invention relates to a method for the acoustic analysis of a machine (M), comprising the acquisition of at least one acoustic signal supplied by at least one microphone (7) positioned in the machine, characterised in that it comprises the following steps: separation of at least one acoustic signal into a plurality of sound sources, said signal being modelled as a mixture of components, each one corresponding to a sound source; for at least one separate sound source, determination of a characteristic acoustic signature; comparison of at least one characteristic acoustic signature with at least one reference acoustic signature recorded in a reference database (5).
G01M 15/12 - Testing internal-combustion engines by monitoring vibrations
G01N 29/14 - Investigating or analysing materials by the use of ultrasonic, sonic or infrasonic waves; Visualisation of the interior of objects by transmitting ultrasonic or sonic waves through the object using acoustic emission techniques
G01N 29/44 - Processing the detected response signal
G01N 29/46 - Processing the detected response signal by spectral analysis, e.g. Fourier analysis
55.
DEVICE FOR TRANSFERRING OIL BETWEEN TWO REPOSITORIES ROTATING RELATIVE TO EACH OTHER, AND PROPELLER TURBOMACHINE FOR AN AIRCRAFT WITH SUCH A DEVICE
The device (20) comprises two outer and inner concentric rings (22, 23), one of which is connected to an oil supply from one of the repositories, the other ring being connected to the other repository, the oil flowing between said rings, and bearings between the rings in order to change repositories between the two rings. According to the invention, the device (20) further comprises a flexible means (31) forming a shock absorber, provided between a first of said rings and an intermediate ring (41) that is separated from a second of said rings by said bearings (25), said flexible means (31) defining a deformable sealed chamber (32) in which oil travels between the two repositories.
The invention relates to a method and system for monitoring an aircraft engine (2), comprising: - acquisition and processing means (11) configured to collect a time signal for a residual temperature margin at the exhaust gas outlet from said aircraft engine (2), - acquisition and processing means (11) configured to smooth said time signal, thus forming a curve representative of said residual temperature margin, - acquisition and processing means (11) configured to identify descending pieces of said first curve, - acquisition and processing means (11) configured to build a second curve by concatenating said descending pieces, said second curve being continuous while being restricted to said descending pieces of said first curve, - acquisition and processing means (11) configured to build a prediction model from said second curve to determine at least one failure prognosis indicator.
In order to improve the cooling of an annular wall (13) of a turbomachine combustion chamber provided with microperforations (53) and, in particular, the cooling of a region of the wall facing a wake (52) caused by an ignition plug, deflector means (60, 68) are proposed, these being designed to deflect the air (34') bathing the ignition plug towards a mid plane (P) of the wake (52) and in the direction of the annular wall (13) of the combustion chamber so as to increase the pressure of the air within the wake (52) near the annular wall (13).
Compacting assembly comprising a shaping mould (24) delimiting a housing open at the top able to receive a precut woven preform (10a), and a compacting tool (128) that is able to move vertically and forms, with the shaping mould (24), a compacting assembly for compacting said preform placed beforehand in the housing. The compacting tool (128) comprises at least one root portion (128A). Application to the manufacture of turbomachine composite blades.
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
A fibrous structure (200) comprises a preform portion (210) formed as a single piece by three-dimensional weaving between a first plurality of layers of threads and a second plurality of layers of threads, the preform portion corresponding to all or part of a fibrous reinforcement preform for a component made of composite. The fibrous structure (200) comprises, outside of the preform portion (210), one or more layers of two-dimensional woven fabric (220a, 220b), each layer of two-dimensional woven fabric grouping together the threads (2010a) of one same layer (201a) belonging at least to the first plurality of layers of threads and situated outside of the preform portion (210).
D03D 25/00 - Woven fabrics not otherwise provided for
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
The invention relates to a fuel injector (10) such as an injector for an annular combustion chamber of a turbomachine, comprising a downstream head (16) having a central outlet (22) and an annular peripheral outlet (24) surrounding the central outlet (22), and an injector arm (12) upstream of the head (16) comprising a coaxial central channel (18) and a coaxial annular channel (20), characterised in that the central channel (18) is in fluid communication with the peripheral outlet (24) and the annular channel (20) is in fluid communication with the central outlet (22).
The invention relates to a combustion chamber for a turbine engine, including an annular bottom wall (18) provided with injection systems (20) each centred on a respective axis (24) and each having an upstream end forming a socket (26') intended for receiving a head of a fuel injector, and an annular fairing (40') covering said bottom wall (18) and including injector-passage openings (42) arranged respectively opposite said injection systems (20), wherein said annular fairing (40) comprises air-intake openings separated from said injector-passage openings (42), and said socket (26') of each injection system passes through the corresponding injector-passage opening (42) and includes, at the upstream end thereof, a flange (62) having a free end (64) separated from said axis (24) of the injection system by a first distance (d1) which is greater than a second distance (d2) separating an edge of said injector-passage opening and said axis.
The invention relates to a compacting assembly comprising a forming mould (24) defining an upwardly open housing that can receive a previously cut woven preform (10a), and a vertically mobile compacting tool (128), and forming, with the forming mould (24), an assembly for compacting said preform previously placed in the housing. The compacting tool (128) comprises at least one foot portion (128A). The compacting tool comprises at least three separate compacting blocks (1281-1287). The invention is applicable to the production of composite fan blades for turbomachines.
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29C 43/36 - Moulds for making articles of definite length, i.e. discrete articles
The invention concerns a rotary assembly for a turbomachine, comprising a disk of which the outer periphery is formed from an alternation of cavities and teeth (12), and blades extending radially from the disk and of which the roots (16) are engaged axially and held radially in the cavities of the disk. According to the invention, the teeth of the disk and the blade roots comprise, at the upstream and/or downstream axial ends of same, axial shoulders (74, 76) disposed circumferentially end-to-end in alternation and together forming a cylindrical surface (78) facing radially towards the inside of the disk.
The present invention concerns a hydraulic ram comprising a fixed support (30), a cylinder (24) that is movable in translation relative to the support, a piston (22) secured inside the cylinder delimiting two chambers with the cylinder (24) and a device for supplying the chambers with hydraulic fluid upstream from the fixed support (30, 31). The ram is characterised by the fact that the supply device comprises telescopic channels (25, 26, 28), each telescopic channel comprising two tubular elements sliding one into the other, a first tubular element being rigidly connected to the fixed support (31) at one end and the second tubular element being rigidly connected to the cylinder at at least two points separated from each other along a generatrix of the cylinder. The invention is applicable to controlling the pitch of the blades of a turbine engine propeller.
B64C 11/38 - Blade pitch-changing mechanisms fluid, e.g. hydraulic
B64C 11/48 - Units of two or more coaxial propellers
F15B 15/14 - Fluid-actuated devices for displacing a member from one position to another; Gearing associated therewith characterised by the construction of the motor unit of the straight-cylinder type
65.
METHOD FOR MONITORING THE ENERGY DENSITY OF A LASER BEAM BY IMAGE ANALYSIS AND CORRESPONDING DEVICE
The invention relates to a method for monitoring (S) the energy density of a laser beam (3) using parameters of the laser beam (3) including the following steps: regularly applying the laser beam (3) to a reference substrate and measuring (S4), with each application, the resulting light intensity; identifying (S6, S7, S8) a change in the light intensity on the reference substrate between at least two measurements; and, when the change in the light intensity is higher than a predetermined threshold, determining the unstable parameter or parameters of the energy density of the laser beam (3) (S8).
B22F 3/105 - Sintering only by using electric current, laser radiation or plasma
B01D 53/00 - Separation of gases or vapours; Recovering vapours of volatile solvents from gases; Chemical or biological purification of waste gases, e.g. engine exhaust gases, smoke, fumes, flue gases or aerosols
B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
B22F 5/04 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product of turbine blades
66.
AFTERBODY FOR A TURBOJET ENGINE COMPRISING A NOZZLE PROVIDED WITH A THRUST REVERSER SYSTEM THAT INCORPORATES A CROWN OF NOISE-REDUCING CHEVRONS
The invention concerns an afterbody for a turbojet engine having a central axis (XX'), provided with a nozzle (1) comprising two doors (3, 4) mounted facing each other between two lateral beams (5, 6) pivoting around axes (15, 16, 17, 18) defining a pivot direction, between a retracted position, in which a middle portion (19a, 20a) of the downstream edge of said doors (3, 4) forms the edge (24) of the outlet section (S) of the nozzle (1) combined with the downstream edges (11, 14) of the two lateral beams (5, 6), and a deployed position, in which said middle portions (19a, 20a) of the downstream edges of the pivoting doors (3, 4) come together so as to block the channel between the two lateral beams (5, 6) in order to reverse the thrust of the turbojet engine gases, the edge of the outlet section of the nozzle further having a crown of noise-reducing chevrons (26) alternating with indentations (25) and the afterbody being characterised in that each noise-reducing chevron (26) formed on the middle portion (19a, 20a) of the downstream edge of a pivoting door is opposite an indentation in the middle portion (20a, 19a) of the other pivoting door (4, 3) in the direction perpendicular to said pivot direction, the shapes of the chevrons (26) and of the indentations (25) being defined to match in such a way that each chevron (26A) fits into the opposing indentation (25A) when the pivoting doors (4, 3) are deployed.
The invention proposes an arrangement of a tank of lubrication liquid for an aircraft turbomachine (12), under a nacelle cowling (10) covering said turbomachine (12), with the tank being arranged in the volume located between the turbomachine (12) and the nacelle cowling (10), with the nacelle cowling (10) comprising an opening (22) for access to said volume, characterised in that the tank (18) is fastened to an internal face (20a) of a panel (20) closing off said opening (22), with the unit formed by the panel (20) and the tank being added onto the nacelle cowling (10) and able to be mounted and dismounted from the nacelle cowling (10).
Case structure interposed between the engine and the nacelle of an aircraft, comprising: a shell ring surrounding the engine having an engine axis and comprising a fixed part (30) and a plurality of sectors (12A, 12B), at least one radial arm (14) for connecting to the nacelle, a plurality of fastening means for securing this plurality of shell ring sectors together or to at least one radial arm, the plurality of fastening means comprising two series of holes made in two parallel side walls of the shell ring sectors and intended respectively to receive two series of screws (28A, 28B) that each pass through a corresponding set of orifices in an adjacent shell ring sector or in an adjacent radial arm, an articulation (24A, 24B) furthermore being disposed between each of the shell ring sectors and a fixed part of the shell ring so as to allow these shell ring sectors to pivot individually about a common rotation axis perpendicular to these side walls disposed along the engine axis.
The invention concerns a fuel system (1) for a turbomachine comprising: - a control circuit (9a), - a main circuit (9b), - a flow rate regulator (2), suitable for regulating the flow of fuel in the control circuit (9a) and in the main circuit (9b) depending on the speed of the turbomachine, and - a drain tank (4) being designed to draw fuel from, store fuel in and drain fuel into the main circuit (9b) on the basis of the pressure difference between the main circuit (9b) and the tank (5) or the high-pressure pump (6) to which it is connected.
Case structure interposed between the engine and the nacelle of an aircraft, comprising: a shell ring surrounding the engine and comprising a plurality of sectors (12A, 12B), at least one radial arm (14) for connecting to the nacelle, a plurality of fastening means for securing this plurality of shell ring sectors together or to the at least one radial arm, the plurality of fastening means comprising a plurality of connecting tabs (20) that are each provided with two parallel series of holes that are intended to respectively receive two series of cap screws that pass, in the case of a first series, through a corresponding set of orifices in a shell ring sector and, in the case of a second series, through a corresponding set of orifices in an adjacent shell ring sector or in an adjacent radial arm.
Case structure of an aircraft, comprising: a shell ring surrounding the engine and comprising a plurality of sectors (12A, 12B), a number of radial arms (14), each mounted between two adjacent shell ring sectors and each having a base (14A), a number of means for fastening these shell ring sectors to the bases of the radial arms, these fastening means having a number of mounting plates (20) that are each mounted in a flush manner in grooves (22) with a shape corresponding to the shell ring sectors and a set of cap screws (26A-26D) that are held captive by the mounting plate as a result of a difference in diameter between the body and the thread of said screws, a set of captive nuts being fixed to the bases so as to ensure, by each receiving the thread of a captive screw of this set of cap screws, a secure connection between the shell ring sectors and the radial arms.
A rotating assembly comprising a transmission member and an oil distribution system for supplying oil to the transmission member in order to lubricate same. According to the invention, the oil distribution system (50) comprises at least one oil transfer chamber (52) provided with at least one supply port (53) designed to receive oil from the outside of the rotating assembly; the transmission member (3) comprises at least one rotating part (35) provided with at least one oil receiving chamber (37); at least one connecting pipe (70) fluidly connects the oil transfer chamber (52) and the oil receiving chamber (37); the oil distribution system (50) is rigidly connected to said rotating part (35) of the transmission member (3) for rotation therewith; and the rotating assembly is configured so as to allow a given relative axial and/or radial movement between said rotating part (35) of the transmission member (3) and the oil distribution system (50).
The invention relates to a bypass turbojet comprising an exhaust housing having a central hub (13) and connecting means (11) that can transmit the forces generated by the turbojet to the structure of the aircraft that it propels, said connection means being two arms extending radially from the central hub in order to cross the cold flow of said turbojet and being characterised in that they are secured to said central hub and positioned in a diametrically opposed manner in relation to each other. An additional connection means (14) extends between the hub (13) and the area (9) for securing the exhaust housing to the structure (12) of the aircraft in order to transmit the exceptional dimensioning loads, said connection means being on standby during normal use, without any transmission of force between said hub and said area.
F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
B64D 27/26 - Aircraft characterised by construction of power-plant mounting
F02C 7/20 - Mounting or supporting of plant; Accommodating heat expansion or creep
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
74.
METHOD FOR THE HIGH-TEMPERATURE SHAPING OF A METAL BLADE REINFORCEMENT
A shaping method implementing a shaping tool suitable for the high-temperature shaping of a preformed metal part having two lateral fins (32, 34) extending from a nose (36), comprising installing said preformed metal part in a first lower die of the tool, holding the preformed metal part in a first predefined position by means of a first movable central insert, shaping one of the lateral fins of the preformed metal part into the final shape of same in alignment with the nose by moving, by means of a first movable upper die, turning the preformed metal part, installing the preformed metal part in a second lower die (52) of said tool, holding the preformed metal part in a second predefined position by means of a second movable central insert (54), and shaping the other lateral fin of the preformed metal part into the final shape of same in alignment with the nose by moving, by means of a second movable upper die (56).
B21K 3/04 - Making engine or like machine parts not covered by ; Making propellers or the like blades, e.g. for turbines; Upsetting of blade roots
B23P 15/04 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from several pieces
75.
METHOD AND SYSTEM FOR INJECTING FUEL INTO AN ENGINE COMBUSTION CHAMBER
The invention relates to a system for injecting fuel into a combustion chamber of an engine, comprising at least two fuel circuits, one with a constant flow rate (1) and the other with an intermittent flow rate (2), members (4, 5) for dosing and distributing fuel between the two circuits and a control means (3) for controlling said members, the intermittent flow circuit being able to be bled, characterised in that, upon receiving an order to fill the circuits with fuel after the intermittent flow circuit has been bled, the control means (3) is suitable for - controlling the dosing and distribution members (4, 5, 6, 7) to obtain a predetermined fuel flow rate greater than the flow rate corresponding to the filling order and providing the resulting fuel surplus to the intermittent flow circuit, during a predetermined duration.
The invention relates to a method for assembling two blades (21, 23) of a turbomachine nozzle, comprising: positioning of a first surface (20) of a first blade (21) and a second surface (22) of a second blade (23) facing one another, said first (20) and second (22) surfaces being spaced apart from one another by an assembly clearance (E); and the vapour-phase aluminising of the first (20) and second (22) surfaces, such as to fill the assembly clearance (E).
The invention relates to a turbine engine stator blade (34) comprising a plurality of blade sections (35) stacked along a radial axis Z. On a bottom portion of the blade, from 0 to 50% of the total height thereof, the leading edge (BA) of each section is in front of the leading edge of the section corresponding to the total height, and the maximum thicknesses are measured at positions spaced apart from the leading edge (BA) of at least 50% of the chord.
The invention relates to an assembly which includes: two rotationally symmetrical components (10, 20) for supporting the blades of a turbine engine, arranged one inside the other concentrically about a turbine engine axis, and a system (30) for locking the components (10, 20) such as to prevent the relative translation of same in the axial and radial directions relative to said axis, the system including a slotted ring (40) comprising a U-shaped cross-section suitable for receiving one end of the components (10, 20), the assembly being characterised in that the ring (40) and one of the components (10, 20) are shaped such as to allow the ring to be interlocked on the component, and in that the locking system (30) also includes a member for stopping the rotation of the ring relative to the component with which the latter is interlocked. The invention further relates to a method for assembling such an assembly.
The epicyclic reduction gear (10) comprises a planetary input shaft (35), planet gears (36) meshing around said shaft and supported by a planet carrier (37), and two transverse sides (41, 43). According to the invention, the reduction gear comprises at least one fluid transfer pipe (54) suitable for being connected to a fluid supply source and passing through the planet carrier (38), being linked in rotation to same, from a first of said transverse sides (41) of the reduction gear to a second of said transverse sides (43) of the latter in order to open on the outside of the reduction gear and dispense the fluid.
This module (5) allows the filtering of a raw setpoint (N1_CMD_OP) intended for a corrector network (6) in an engine regulating system (20). It comprises: - a module (54) for detecting a filtering condition of said raw setpoint, and; - means for providing said corrector network with a filtered setpoint (N1_CMD_LIM) in place of said raw setpoint, when the filtering condition is detected.
F02C 9/28 - Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
G05B 19/404 - Numerical control (NC), i.e. automatically operating machines, in particular machine tools, e.g. in a manufacturing environment, so as to execute positioning, movement or co-ordinated operations by means of programme data in numerical form characterised by control arrangements for compensation, e.g. for backlash, overshoot, tool offset, tool wear, temperature, machine construction errors, load, inertia
G05B 19/416 - Numerical control (NC), i.e. automatically operating machines, in particular machine tools, e.g. in a manufacturing environment, so as to execute positioning, movement or co-ordinated operations by means of programme data in numerical form characterised by control of velocity, acceleration or deceleration
81.
METHOD FOR MONITORING A VALVE OF AN AIRCRAFT ENGINE
The invention relates to a method and system for detecting first signs of failure in an aircraft engine valve, comprising: acquisition means configured to acquire output pressure measurements of said valve (3) and command and context data related to said valve (3); processing means (23) configured to define a series of indicators of the first signs of failure in accordance with said output pressure measurements and said command and context data; processing means (23) configured to monitor changes over time in each indicator of said set of indicators of the first signs of failure; and processing means (23) configured to detect a possible deviation of at least one indicator among said set of indicators, said deviation representing the first signs of failure of said valve.
F16K 37/00 - Special means in or on valves or other cut-off apparatus for indicating or recording operation thereof, or for enabling an alarm to be given
A mobile turbomachine blade comprising a vane defined by a stack of elementary sections that are planar in a radial direction, each elementary section being positioned radially at a height H, height H being expressed as a percentage of the total height of the vane, and being located by the sweep angle ? of same and the dihedral angle v of same, in which the change in sweep angle ? depending on height H is such that sweep angle ? reaches a maximum for a height ???, value ??? being between 5% and 40%, and sweep angle ? is increasing between 0% and ???, and in which the change in dihedral angle v depending on height H is such that dihedral angle v is a decreasing function of height H, for heights H between 0% and a value Hvl, value Hvl being between 10% and 40%.
The invention proposes a casing (30) of an aircraft turbine (16), intended to support a set of ring sectors (28) which partly delimits a channel for the passage of a gas stream through the turbine (16), where the casing (30) includes means for dynamic adjustment of the radial position of the ring sectors (28) by controlled injection of an air stream on to portions (36) of an annular wall (34) of the casing (30), where the casing (30) includes an upstream radial tab (38) which connects an upstream end of each ring sector (28), in the flow direction of the gas stream, to the casing (30), and a downstream radial tab (40) which connects a downstream end of each ring sector (28) to the casing (30), where both upstream and downstream radial tabs (38, 40) are made from a single piece with the casing (30), characterised in that each radial tab (38, 40) is made as two portions (42, 44) from different materials.
The invention relates to a method for manufacturing a turbine engine casing (1), characterised in that said method includes the steps of: manufacturing (E1) a plurality of sectors (2), at least one portion of the sectors (2) being manufactured by casting and including, on the surface thereof, fastening elements (3) produced during the casting step, assembly bands (8) being produced at the ends of the sectors (2) during the step of manufacturing the sectors (2) by casting, by means of which the sectors (2) can be assembled; and assembling (E2) the sectors (2) end-to-end such as to form a ring (5) of the casing (1). The invention also relates to a turbine engine casing (1).
The present invention concerns a method for modelling at least a part of a blade (2) of a non-streamlined propeller (1), the part of the blade (2) having an offset (3), the method being characterised in that it comprises implementing, using data processing means (11) of a piece of equipment (10), steps of: (a) Parameterisation of at least one Bézier curve representing a deformation of said blade (2) characterising the offset (3), defined by: a. First and second end control points (PCU1, PCUK); b. At least one intermediate control point (PCU,, i? [[2,k-1]] disposed between the end points (PCU1, PCUK), the parameterisation being implemented according to at least one deformation parameter and said cutting height in the blade (2), on the basis of which the abscissa of the intermediate control point (PCUi,) and the ordinate of the second end point (PCUK) are expressed; (b) Determination of optimised values of the deformation parameter or parameters; (c) Outputting the values determined in this way to an interface (13) of said piece of equipment (10).
METHOD OF ESTIMATION ON A CURVE OF A RELEVANT POINT FOR THE DETECTION OF AN ANOMALY OF A MOTOR AND DATA PROCESSING SYSTEM FOR THE IMPLEMENTATION THEREOF
The invention relates to a method of estimation on a curve of a relevant point for a detection of an anomaly of a motor, said curve representing an evolution as a function of time of physical operating parameters of the motor measured by sensors on said motor, implemented by a computer linked to first means of storage, said first means of storage storing at least one profile comprising a binary code, each component of which codes a direction of variation between two consecutive characteristic points of at least one learning curve, a model making it possible to estimate a relevant point on the basis of a set of characteristic points of a curve and a filter, said method comprising: a/ (F1) the selecting of a profile stored in the first means of storage; b/ (F2) the application of the filter of the profile selected to said curve; c/ (F3) the determination of a set of characteristic points of said filtered curve and of a binary code, each component of which codes the direction of variation of two consecutive characteristic points belonging to said set of characteristic points; d/ (F4) the comparing of the determined code and of the code of the profile selected; e/ (F5) as a function of said comparison, the estimating of the relevant point on said curve on the basis of the characteristic points of said filtered curve and of the model of the profile selected.
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups ; Air intakes for jet-propulsion plants
87.
COVER OF A TURBOMACHINE CENTRIFUGAL COMPRESSOR CAPABLE OF BEING RIGIDLY CONNECTED VIA THE DOWNSTREAM SIDE NEAR TO THE UPSTREAM EDGE OF SAME, AND TURBOMACHINE COMPRISING THIS COVER
The invention concerns a cover (1) of a centrifugal compressor intended to be rigidly connected to a turbomachine housing (13, 15), which comprises a plurality of ports (16). The cover (1), which further comprises means for fastening same to the housing, is characterised in that a portion of said fastening means (20) is located upstream relative to the ports (16) and can be accessed by a fastening tool (22) through at least one of said ports (16) of the cover (1). The invention also concerns the turbomachine using this cover to form a sealed space (17), in particular with a view to collecting air.
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
F01D 5/00 - Blades; Blade-carrying members; Heating, heat-insulating, cooling, or antivibration means on the blades or the members
F01D 11/14 - Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F04D 27/00 - Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
F04D 29/62 - Mounting; Assembling; Disassembling of radial or helico-centrifugal pumps
88.
DEVICE FOR PROTECTING AGAINST OIL LEAKS TOWARDS THE ROTORS OF A TURBOMACHINE TURBINE
The invention concerns a device for protecting against oil leaks towards the rotors of a turbomachine turbine (2), comprising an upstream cavity (18) in which an oil/air mixture can circulate, said upstream cavity being delimited downstream by a journal (8) of the turbine rotor and on the outside by a sealing flange (14) of the journal extending axially in an upstream direction, a downstream cavity (20) opening towards the discs (12) of the turbine and communicating with the upstream cavity via a plurality of ventilation holes (22) provided in the journal, and an upstream annular deflector (24) fixed to the sealing flange of the journal and extending radially inwards towards the ventilation holes, the upstream deflector having oil passage ports (26) provided at the end of same secured to the sealing flange of the journal.
The invention relates to a method for assembling a first turbomachine part with at least one second turbomachine part. The assembly method comprises the following steps: injection of a vulcanisable elastomer, preferably a silicone (145) that can be vulcanised at room temperature, known as RTV silicone, in an injection zone (140) at the junction between the first and second part; and local heating of the injection zone (140) such as to vulcanise the vulcanisable elastomer. The invention also relates to an assembly (20) used during such an assembly method.
B29C 45/16 - Making multilayered or multicoloured articles
B29C 45/14 - Injection moulding, i.e. forcing the required volume of moulding material through a nozzle into a closed mould; Apparatus therefor incorporating preformed parts or layers, e.g. injection moulding around inserts or for coating articles
B29C 65/70 - Joining of preformed parts; Apparatus therefor by moulding
The invention relates to an aircraft propeller blade including a structure having an airfoil, consisting of at least one fibrous reinforcement (200) obtained by three-dimensional weaving of threads and made denser by a matrix and a spar (60) including a bulged portion (62) extending outside of the fibrous reinforcement and forming the shank of the blade, and a shaped portion (61) in a recess (206) provided within the fibrous reinforcement (200). The fibrous reinforcement (200) comprises a disconnection area (204) forming the recess (206) inside the fibrous reinforcement (200). The disconnection area (204) leads to the lower portion (220) and the rear edge (230) of the fibrous reinforcement (200) such as to form an opening (231) for inserting the shaped portion (61) of the spar (60) into the recess (206) of the fibrous reinforcement. The opening (231) on the rear edge (230) of the fibrous reinforcement (200) extends over a height (H231) that is shorter than the height (H204) of the recess such as to provide, on said rear edge (230), a retaining portion (232) that is at least partially in contact with the shaped portion (61) of the spar (60).
The invention relates to a device (11) for cooling oil for a turbine engine, such as an aircraft turbofan or turboprop, characterised in that it comprises a duct (12) for circulating a flow of cold air (F1); means for injecting oil (16) into the duct; and means (19) for extracting the oil mixed with the flow of cold air (F1), located in the duct (12), downstream from the injection means (16).
The present invention relates to a method for modeling at least part of a non-axisymmetric surface (S) of one portion (2) of a part (1), the portion (2) being bounded by a first and a second extremal planes (PS, PR). The method is characterized in that it includes the implementation, by a means (11) for processing data from a device (10), of the following steps: (a) modeling at least two basic non-axisymmetric surfaces (Sk, SI, SE) making up said surface (S), each of the basic surfaces (Sk, SI, SE) extending between both extremal planes (PS, PR) of the portion (S) that are juxtaposed so as to have a junction of at least C1 at the extremal planes (PS, PR) and at least one discontinuity; (b) constructing at least one C1 connection curve (CR) tangential to the first and second surfaces; (c) locally modifying the first and second basic surfaces (Sk, SI, SE) such as to keep to said connection curve (CR) near said discontinuity; and (d) restoring, on one interface (13) of said device (10), said at least part of the obtained surface (S).
The invention relates to a device (100; 500) for inspecting the surface of an electrically conductive part, including a plurality of Eddy current probes (330, 331, 332...) arranged on a convex surface of the device, and means for applying the probes against the surface to be inspected into which the device is inserted, characterized in that the probes (330, 331, 332...) are attached onto flexible blades (320, 321, 322...) each extending side by side in a longitudinal direction (L) of the device, wherein said application means include a deformable material (490), the compression of which, in the longitudinal direction (L), results in an expansion that is transverse to the longitudinal direction, said expansion deforming said blades (320, 321, 322...) such as to apply the probes (330, 331, 332...) against the surface.
G01N 27/904 - Investigating or analysing materials by the use of electric, electrochemical, or magnetic means by investigating magnetic variables for investigating the presence of flaws using eddy currents with two or more sensors
94.
PROCESS FOR PREPARING FOR REMOVAL A METAL COATING BY THERMAL SPRAYING ON A SUBSTRATE
The invention relates to a process for preparing for removal a metal coating by thermal spraying on a substrate, comprising the following steps: a) depositing an adhesive layer on the zone to be coated with a uniform thickness larger than 10 µm and smaller than 100 µm; b) cold spraying, before the adhesive has dried, a metal powder on the zone to be coated, the powder particles thereby becoming embedded at least partially in said adhesive layer; and c) drying the adhesive, said powder particles remaining imprisoned therein, thereby forming an underlayer suitable for receiving a metal coating deposited by thermal spraying. Application to the protection of the leading edge of blower blades.
The present disclosure concerns a remote connection system capable of being incorporated into an aircraft (1A, 1B, 1C) comprising at least one motor propeller (50A, 50B, 50C) having a plurality of blades (52A, 52B, 52C) capable of rotating relative to a fixed module (10A, 10B, 10C) of the aircraft about an drive shaft (X). The system comprises: an optical transmission device configured such that, when the system is incorporated into the aircraft (1A, 1B, 1C), it transmits a light beam that emerges out of the propeller (50A, 50B, 50C), from at least one transmitting surface (54A, 54B, 54C) of said propeller (50A, 50B, 50C); and an optical device for detecting the light beam, comprising at least one detecting surface (14A, 14B, 14C) that is sensitive to the light beam, and that is capable of being incorporated into the fixed module (10A, 10B, 10C) in such a way that said at least one transmitting surface (54A, 54B, 54C) and said at least one detecting surface (14A, 14B, 14C) repeatedly come to face each other, at a distance from each other, when the propeller (50A, 50B, 50C) rotates relative to the fixed module (10A, 10B, 10C). The present disclosure also concerns an assembly comprising the propeller, the fixed module and the connection system.
The invention relates to a pivot (30) for an air propeller blade, comprising a rotary support (36) which is to be radially mounted on a propeller hub while being able to pivot about a pivoting axis (Z), a blade support (38) comprising a housing (40) for receiving a blade root, and at least one arm (34a, 34b) extending laterally in relation to the pivoting axis and carrying a balance weight (32) forming a counterweight, the arm and the balance weight having a general geometric form of a spherical cap.
The method for resetting a basic digital model (5) according to the invention comprises: - a step (E10) of detecting a stable state of at least one first parameter (T25) of said model, said first parameter being representative of a signal delivered by a sensor (3); - a step (E60) of obtaining a resetting parameter (GainF) of said model, during said stable state of said first parameter (T25), on the basis of said first parameter, a second parameter (PCN12R) of said model, and said basic digital model (5); and - a step (E70) of obtaining a reset model from the basic digital model (5) and the resetting parameter (GainF).
The invention relates to a turbomachine intermediate casing (10) comprising: - an internal hub (11) designed to accept a turbomachine compressor shaft, - an outer shell ring, and - a plurality of radial arms (13) connecting the hub and the shell ring, one of the arms comprising an internal housing of radial shaft (20) for driving auxiliary machines, the casing being characterized in that it further comprises a bevel gearbox (30) for the radial shaft (20) arranged in the outer shell ring, said gearbox being formed as one with the outer shell ring, and the bevel gearbox (30) comprises a housing for gears (31) between a radial shaft and an accessories gearbox, said housing (31) comprising an opening (35) opening to the periphery of the outer shell ring (12). The invention also relates to a turbomachine accessories gearbox drive assembly comprising such a casing.
The invention relates to a device for fixing two parts together, a first part having a fixing flange (2) comprising at least one through-orifice (3) for a threaded shank (4), carried by or integral with a second part (5), and means for making the nut (1) captive with respect to the flange (2) before it is screwed onto the threaded shank (4), characterized in that these means are of the type having a thread, the nut (1) having an externally threaded portion (11, 18) intended to be screwed into an internally tapped portion (8, 15) of the orifice (3) in the flange (2) or of an element (14) attached to the flange (2) and to pass completely through this internally tapped portion (8, 15) until it is housed in a cavity (9, 19) for axially retaining the nut (1), this allowing the nut (1) to rotate freely before it is screwed onto the threaded shank (4). The invention also relates to a turbomachine that utilizes this device and to an associated fixing method.
F16B 5/02 - Joining sheets or plates to one another or to strips or bars parallel to them by means of fastening members using screw-thread
B64D 29/06 - Attaching of nacelles, fairings, or cowlings
F16B 41/00 - Measures against loss of bolts, nuts, or pins; Measures against unauthorised operation of bolts, nuts, or pins
100.
METHODS OF CREATING A DATABASE AND OF FORMULATING A MAP OF OPERATING STATES OF AIRCRAFT AND A METHOD OF MONITORING THE OPERATION OF AN ASSOCIATED AIRCRAFT
The invention relates to a method of creating a database of operating states of an aircraft fleet, in which a computer executes the steps of: - (301) acquiring parameters characterizing a state of operation of an aircraft of said aircraft fleet; - (302) acquiring at least one item of degradation information indicative of a level of degradation of at least one aircraft from among said fleet of aircraft; - (303) storing series of parameters in said database, each series of parameters comprising parameters characterizing a state of operation of one of said aircraft and if said item of degradation information has been acquired, at least one item of degradation information indicative of a level of degradation of the aircraft.
patents
