Abstract: Provided is a method for optimising memory writing in a device implementing a cryptography module and a client module calling functions implemented by the cryptography module. The device includes a random access memory including a first memory zone that is secured and dedicated to the cryptography module and a second memory zone dedicated to the client module. When the client module calls a series of functions implemented by the cryptography module including a first function and at least one second function, with each second function executed following the first function or from a further second function and providing a runtime result added to a runtime result of the preceding series function, each runtime result is added to a value contained in a buffer memory allocated in the first memory. The buffer memory value is copied to the second memory zone following the execution of the last function of the series.
Type:
Grant
Filed:
April 26, 2017
Date of Patent:
October 29, 2019
Assignee:
SAFRAN IDENTITY & SECURITY
Inventors:
Guillaume Dabosville, Philippe Gislard, Victor Servant
Abstract: An outlet guide vane wheel includes guide vanes made of polymer matrix composite material reinforced by fibers, each having a vane root and a vane tip. The vane roots are fastened on a hub of the wheel by a first connection, and the vane tips are fastened on an outer shroud of the wheel by a second connection. The first connection includes a bearing plane secured to the hub and a first backing plate for securing to the hub, with the vane roots being sandwiched between the bearing plane and the first backing plate. The second connection includes a second backing plate for securing to the shroud, with the vane tips being sandwiched between the shroud and the second backing plate.
Abstract: A stator of an aircraft turbine engine, comprising an annular row of fixed vanes and an annular row of arms, wherein the trailing edges of the fixed vanes are positioned substantially in a first transverse plane that is positioned downstream of a second transverse plane that passes substantially through the leading edges of the arms.
Type:
Grant
Filed:
September 29, 2015
Date of Patent:
October 29, 2019
Assignee:
SAFRAN AIRCRAFT ENGINES
Inventors:
Simon Pierre Claude Charbonnier, Matthieu Yoann Perrier
Abstract: When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.
Type:
Grant
Filed:
December 29, 2016
Date of Patent:
October 29, 2019
Assignee:
Safran Aircraft Engines
Inventors:
Thomas Michel Mervant, Erwan Daniel Botrel, Maxime Didier Delabriere, Jean-Armand Marc Destouches
Abstract: An airborne shootings detection and piloting aid equipment is disclosed for an aircraft including a multi-sensor system supplying signals representing an environment of the aircraft including at least one system supplying a signal, referred to as an LWIR/MWIR signal, representing infrared radiation lying the far infrared or mid-infrared range; piloting aid means for obtaining first piloting information including information representing obstacles present in the environment of the aircraft; and shootings detection means for obtaining second piloting information including information representing shootings in the vicinity of the aircraft, the piloting aid means and the shootings detection means using signals representing an environment of the aircraft including at least the LWIR/MWIR signal; and obtaining means for obtaining piloting aid parameters from first and second piloting information.
Abstract: A turboshaft engine includes a casing in which is arranged a gas generator and a free turbine fitted to a power shaft. The power shaft is configured to be mechanically connected to/disconnected from a reduction gearbox. The turboshaft engine includes at least one centralizer movable between an active position, in which the centralizer forms a bearing for the power shaft and which corresponds to a mechanical disconnection between the power shaft and the reduction gearbox, and a passive position, in which the centralizer is distanced from the power shaft and which corresponds to a mechanical connection between the power shaft the reduction gearbox.
Abstract: A thrust reverser for a nacelle of an aircraft turbojet engine is provided that includes a thrust reverser cowl movable along a direction parallel to a longitudinal axis of the nacelle and a variable section outlet nozzle extending from the thrust reverser cowl. The thrust reverser further includes an actuator, a first locking device to lock the thrust reverser cowl, a second locking device to lock the variable section outlet nozzle, and a reset lever. The reset lever is pivotally driven by a locking pin secured to the cowl and pivots from a rest position to a reset position.
Abstract: A method for cutting a preform usable for production of a turbomachine part and including a weaving of a plurality of threads, the threads including single threads that are visually identifiable, along a cutting contour calculated based on a preform model in which the threads have a reference arrangement, the method including: taking an image of the preform; processing the image to determine a deviation in an arrangement of the threads which are visually identifiable relative to the reference layout; correcting the cutting contour according to the deviation; and cutting the preform along the corrected cutting contour.
Type:
Grant
Filed:
June 26, 2018
Date of Patent:
October 29, 2019
Assignee:
SAFRAN AIRCRAFT ENGINES
Inventors:
Yann Marchal, Philippe Marolle, Claire Rousseau
Abstract: When cold and in the non-coated state, a turbine blade includes an aerodynamic profile that is substantially identical to a nominal profile determined by the Cartesian coordinates X, Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.
Type:
Grant
Filed:
December 29, 2016
Date of Patent:
October 29, 2019
Assignee:
SAFRAN AIRCRAFT ENGINES
Inventors:
Pascal Pierre Routier, Vincent Nicolas Leonardon, Daniel Marius Man, Pierre Hervé Marche
Abstract: A method for the non-destructive testing of the heating of a part made from polymer material, the method comprising the following steps: a) carrying out a measurement by infrared spectroscopy on a part to be tested and extracting therefrom at least one of absorbance values and transmittance values according to a spatial frequency; and b) from the measurement of at least one of absorbance and transmittance, determining the period of time during which said region of the part to be tested has been subjected to a given heating temperature and determining said heating temperature, using a reference database comprising at least one of absorbance measurements and transmittance measurements, the measurements established over a plurality of reference samples made from polymer material that have been subjected to a given temperature during a given period of time.
Type:
Grant
Filed:
January 30, 2018
Date of Patent:
October 29, 2019
Assignee:
Safran Aircraft Engines
Inventors:
Jean-Louis Romero, Angélique Melody Marine Alexia Mainczyk, Geoffrey Martin
Abstract: The shank of a blade includes a first side and a second side provided with a depression. By concentrating this depression on a single side, a shank that is more rigid and resistant with an equal lightening is obtained, with the stress concentrations depending on the depth of the depression but being absent from first flat side. The vibrations and bending deformations are also reduced.
Type:
Grant
Filed:
January 16, 2018
Date of Patent:
October 29, 2019
Assignee:
SAFRAN AIRCRAFT ENGINES
Inventors:
Marco Nucci, Damien Merlot, Jacky Rodriguez
Abstract: A brush seal system, for application as an air/oil seal in a turbomachine shaft bearing, of axis of rotation, with the seal including a layer of carbon bristles, held between a first ring disposed upstream of a flow of air that passes through the seal, and a second ring disposed downstream of the flow of air, with a surface of the layer of bristles, which is turned towards the first ring, including a non-metal flexible element that makes it possible to stop a portion of flow of air that flows through the seal in a direction substantially parallel to the axis of rotation.
Abstract: A bowl for supplying oil to at least two oil-distribution circuits which are connected to a planet carrier of an epicyclic reduction gearbox, the planet carrier rotating and the oil coming from a fixed oil ejector, the bowl being designed to be secured to the planet carrier and having a substantially cylindrical shape and being open radially towards the inside with respect to an axis. The bowl is divided into a circumferential succession of separate cups each designed to communicate with one of the oil distribution circuits A reduction gearbox with its supply device and a method of operation in a turbomachine are also disclosed.
Abstract: The invention relates to an injection mold (4) for manufacturing a rotary part made of a composite material having outer flanges, comprising a mandrel (2) on which a fiber reinforcement is intended to be held (9) and including a central wall (6) the profile of which matches that of the part to be manufactured and two side plates (8), molding wedges (12) for bearing against a non-covered surface of the fiber reinforcement, and two outer bells (14) for covering the molding wedges and the plates of the mandrel with sealing O-rings (24) being inserted between the bells and the plates of the mandrel, the bells each being provided with an attachment flange (20) for bearing against one another with at least one sealing O-ring (26) inserted therebetween, the attachment flanges of the bells being clamped against one another.
Abstract: The present application concerns a test engine hood for a turbine engine, such as a double flow turbine. The test hood allows replacement of a flight engine hood during tests on a test bench on the ground where the temperature conditions could damage the flight hood. The test hood includes a tubular wall of carbon-fiber epoxy composite, and metal flanges upstream and downstream. To provide thermal protection, the test hood includes a layer of silicone with a majority of polysiloxane. The layer covers the entire inner surface of the wall to create a barrier. The present application also concerns a method for testing a turbine engine on a test bench, where the turbine engine is fitted with a test casing. The present application also concerns a use of silicone for thermal insulation of the inside of the test hood of the turbine engine on a test bench on the ground.
Type:
Grant
Filed:
September 28, 2015
Date of Patent:
October 29, 2019
Assignee:
SAFRAN AERO BOOSTERS SA
Inventors:
Quac Hung Tran, Pierre Croes, Aurore Noelmans
Abstract: A turbine engine includes two rotary shafts and a lubrication unit. The lubrication unit has at least one pump which a casing inside of which a rotor is mounted and driven by one of the rotary shafts. The pump casing is rotated by the other rotary shaft such that the actuation of the pump depends on the difference between rotational speeds of the shafts. The shafts are a drive shaft and a fan shaft, respectively, wherein the fan shaft is driven by the drive shaft by means of a reduction gear which is lubricated by the lubrication unit. The reduction gear is annular and the pump passes axially therethrough.
Type:
Grant
Filed:
February 5, 2015
Date of Patent:
October 29, 2019
Assignee:
SAFRAN AIRCRAFT ENGINES
Inventors:
Benoît Jean Henri Gomanne, Michel Gilbert Roland Brault, Thomas Chauveau, Bellal Waissi
Abstract: An aircraft landing gear assembly includes a structural member coupled to another part by a joint including engagement formations or a relatively high friction coating arranged such that the joint inhibits pivotal movement of the structural member when the structural member is compressively loaded in an axial manner.
Abstract: The disclosure relates to a turbofan nacelle including a thrust reverser, the thrust reverser including a movable cover moving back from a closed position in which the thrust is not reversed to an open position for uncovering cascades that reverse the direction of the flow of cold air which is diverted from the annular stream of secondary air, the movable cover including a radially outer portion intended for being adjacent to a leading edge of a wing of an aircraft. The movable cover includes on the radially outer portion at least one cut-out intended for avoiding interference with a movable slat of the leading edge of the wing of the aircraft, as well as a panel for closing the cut-out, the closing panel including a stationary portion at least partially covered by an upstream portion of the movable cover when the movable cover is in the open position.
Abstract: The invention relates to an annular combustion chamber in a turbine engine, comprising two coaxial walls of revolution, one an internal wall (12) and the other an external wall (14), between which emerge fuel injectors (19) each engaged in centring means (36) which are moveable in the radial direction in support means (38), the chamber also comprising means (50, 68) of axial retention in the upstream direction of the centring means. According to the invention, the means of axial retention in the upstream direction are removably fixed to at least one of the internal (12) or external (14) walls of revolution.
Type:
Grant
Filed:
December 17, 2014
Date of Patent:
October 29, 2019
Assignee:
SAFRAN AIRCRAFT ENGINES
Inventors:
Brice Le Pannerer, Laurent Bernard Cameriano, Pierre-François Simon Paul Pireyre
Abstract: Hydraulic valves for dampening pressure spikes and associated methods are disclosed herein. In one embodiment, a hydraulic valve for dampening pressure spikes includes: a spool configured to move axially inside the hydraulic valve; and a sleeve configured to at least partially house the spool. A location of the spool with respect to the sleeve may determine a flow of a working fluid through the hydraulic valve. A viscous damper is at least partially housed inside an opening in the spool, and a viscous friction between the viscous damper and the opening in the spool slows a motion of the spool.