The invention refers to an overheat detection system for an aircraft, the system comprising a first bleed monitoring computer, BMC1, configured to identify leakages in a pneumatic system, the BMC1 including a first optical controller, a second bleed monitoring computer, BMC2, the BMC2 including a second optical controller, an optical fiber link connecting the first optical controller of the BMC1 and the second optical controller of the BMC2 for communication between the BMC1 and the BCM2 and between the first optical controller and the second optical controller, wherein the first and the second optical controllers are configured to detect overheat of the optical fiber link based on a wavelength shift of a modulated optical signal transmitted through said optical fiber link, and transmit signals to the first BMC1 and the second BMC2 based at least on the detected overheat.
In order to avoid the intervention of an operator on a direction control device (18) with which an aircraft towing system (4) is equipped, this system comprises: - a first towing sling (10a) comprising a front end (12a) connected to a towing vehicle (2), and a rear end (13a) connected to a first main landing gear (8a); - a second towing sling (10b) comprising a front end (12b) connected to the vehicle, and a rear end (13b) connected to a second main landing gear (8b); and - the direction control device (18) connected to a front landing gear (6), and arranged between the first and second slings, this device comprising a lateral direction control means (24) having a first lateral end (30a) mounted on the first sling (10a), and a second lateral end (3013) mounted on the second sling (10b).
3.
SUPPRESSION OF ELECTROSTATIC DISCHARGE NOISE BY MEANS OF CONDUCTION BETWEEN A TIERED METAL ELEMENT AND THE WIRING SYSTEM OF A GLAZING UNIT
The invention relates to laminated glazing comprising a first glass sheet and a second glass sheet (1; 2) bonded by means of an adhesive interlayer (2), a peripheral zone of the laminated glazing being covered with a tiered metal element (7), said laminated glazing comprising a heating wiring system and/or electrically conductive layer (11) provided with bus bars (12), multiple sensors and other optional electrical elements that are connected to a power supply by means of the connector (13) of the laminated glazing. An electrical conductor (21) connects the tiered metal element (7) to: a bus bar (12) of the heating wiring system and/or electrically conductive layer (11); and/or the enclosure (14) of the connector (13) of the laminated glazing, connected to the mass of the structure for mounting the laminated glazing; and/or a contact (15; 16; 17) of the connector (13). The invention also relates to a method for producing said glazing and to the uses thereof.
B32B 17/10 - Layered products essentially comprising sheet glass, or fibres of glass, slag or the like comprising glass as the main or only constituent of a layer, next to another layer of a specific substance of synthetic resin
The invention relates to a laminated glazing including first and second glass sheets (1; 3) bonded by a first intermediate adhesive layer (2), a peripheral zone of the laminated glazing being covered by a stepped metal element (7), wherein a pane retainer (31) rigidly secured to the mounting structure of the laminated glazing contacts the laminated glazing in order to secure it to its mounting structure, and an electrical conductor (21) connects the stepped metal element (7) and the mounting structure of the laminated glazing via the pane retainer (31); the invention also relates to the uses thereof.
B32B 17/10 - Layered products essentially comprising sheet glass, or fibres of glass, slag or the like comprising glass as the main or only constituent of a layer, next to another layer of a specific substance of synthetic resin
The invention refers to a panel structure (1) for an aircraft comprising at least one composite layer (3, 4), and at least one net-shaped layer (2) attached to the at least one composite layer (3, 4), wherein the net-shaped layer (2) comprises a material suitable to improve the impact resistance of the panel structure (1). The net-shaped layer (2) can be used to attach two composite layers (3, 4). Alternatively, the net-shaped layer (2) can be attached to a surface of one composite layer (3, 4). In this last case, the net-shaped layer (2) may be joined to a laminate sheet material (5), and filled with a foam material (6). The invention provides an impact reinforced panel structure (1) capable of withstanding any impact, such as a blade release, or a bird strike, without substantially modifying the manufacturing process
B64C 1/12 - Construction or attachment of skin panels
B32B 5/02 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by structural features of a layer comprising fibres or filaments
B32B 37/00 - Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding
B64C 3/26 - Construction, shape, or attachment of separate skins, e.g. panels
6.
METHOD AND DEVICE FOR INSPECTING THE DAMAGE ON THE SKIN OF AN AIRCRAFT AFTER A LIGHTNING STRIKE
The invention relates to a method and device for inspecting the structural damage to the skin of an aircraft after a lightning strike, wherein the skin comprises a set of marks caused by the lightning strike, said marks including a visual damage (VD) and a structural damage (SD). The method comprises the following steps: for each mark, measuring the surface area of the paint that has been removed by the lightning strike; comparing the measured surface area with a surface area threshold value, wherein the threshold value relates to the structural damage (SD) of each mark.
The invention relates to a central wing box for an aircraft, comprising a top panel, a bottom panel, a front spar, a rear spar, and at least one secondary rib. According to the invention, a secondary rib comprises two stiffeners (410A, 410B), a first stiffener (410A) extending adjacent to the top panel and to one of the front spar and the rear spar, and a second stiffener (410B) extending adjacent to the bottom panel and to the other of the front spar and the rear spar. The invention employs a reduced number of components, thereby simplifying the construction of the central wing box of an aircraft, particularly the installation of the secondary ribs.
The invention concerns an electro-magnetic wave absorption device (100) to be attached to a wall (12). The absorption device (100) consists of a metal plate (102) to be attached to the wall (12), a dielectric panel (104) remote from the metal plate (102), and delimiting with the metal plate (102) a volume (112) between the metal plate (102) and the dielectric panel (104), a network of resistive dipoles (106), where all resistive dipoles (106) are attached to the same side of the dielectric panel (104), and where each one consists of two metal inserts (108a-b) remote from each other and of a resistance (110) positioned between edges facing two neighbouring metal inserts (108a-b). The absorption device also includes metal walls (114) electrically connected to the metal plate (102) and extending from the metal plate (102) toward the dielectric panel (104).
The object of the invention is a manoeuvre device for at least one aircraft trap door, the said device including a fixed chassis (24) secured to the fuselage (26) of the aircraft, a control handle (22) mobile in rotation and in translation relative to the chassis (24), a mechanism for converting a rotary movement of the said control handle (22) into an opening movement of the said trap. The translation movement enables the control handle (22) to occupy a stowed position in which it is aligned with the exterior surface of the fuselage (26) and a deployed position in which it protrudes relative to the exterior surface of the fuselage (26).
The invention concerns an avionics bay (1) for the installation of at least one electrical module (M), comprising a fluid cooling system and a housing (2) that is complementary to the electrical module (M) and that comprises an open front face (21) through which the module (M) can be removably installed inside said housing (2), and a rear face (20) on which there are arranged electrical connectors suitable for being connected to the electrical module (M), characterised in that the fluid cooling system comprises a cold plate (3) that is disposed on the rear face (20) of the housing (2) that is suitable for cooling the electrical module (M) when said electrical module (M) is installed inside the housing (2).
A fastener (10) of parts (P1, P2) of an assembly comprises a rod (11) formed of a cylindrical drum (12) and an end (13), the rod (11) being a metal part at least partially surface-treated by anodic oxidation (O) and coated on at least a portion of the cylindrical barrel (12) with a lubricating coating (R). Such a fixing element (10) ensures electrical continuity with the parts (P1, P2) during a moderate interference assembly in the assembly.
This elastic device for coupling (7) between an aircraft landing gear wheel (3) and a coaxial member (6) for motorizing said wheel comprises a plurality of arms (9) integral in rotation with said motor drive member from side of a first radially inner end and adapted to be connected to said wheel (3) on the side of their radially outer end, said arms (9) being substantially rigid in a plane perpendicular to the axis of rotation of the wheel and flexible in directions substantially transverse to said plane.
The invention relates to a process for manufacturing a textile preform from a fibrous array (20) with continuous fibres, the process comprising a step for heating of the fibrous array between a matrix (4) and a punch (2) to stiffen this fibrous array. To cut down processing time, the heating step is performed by having a hot gas flow circulate through the matrix (4), the fibrous array (20) and the punch (2).
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]
14.
LIGHTNING CONDUCTOR SYSTEM COMPRISING A LIGHTNING CONDUCTOR STRIP MOUNTED IN AN OFFSET MANNER
The main subject matter of the invention is a lightning conductor system (10) for a support structure (1) having an external face (1a) intended to be subjected to a flow of air and an internal face (1b) opposite the external face (1a), the lightning conductor system (10) comprising an electrically conducting lightning conductor strip (2), intended to be placed on the side of the internal face (1b) of the support structure (1) while being connected to ground, and means (3, 4, 6, 7) of fixing the lightning conductor strip (2) to the support structure (1), the fixing means comprising fixing studs (3) intended to be inserted into orifices formed in the support structure (1), said fixing studs (3) being in electrical contact with the lightning conductor strip (2), the lightning conductor system (10) being characterised in that the lightning conductor strip (2) is mounted in an offset manner with respect to the main axes (X) of the fixing studs (3).
The invention proposes a method for producing a panel comprising an outer skin and stiffeners comprising a connecting portion (18), and at least one protruding portion (16), which comprises a step of draping layers of fibres (24) on at least one insert (30) such that each layer of fibre (24) comprises a connecting portion (28) and a portion (26) protruding downwards relative to the connecting portion (28), a step of draping a layer of fibres (22) forming the outer skin (12), and a step of joining the layers of fibres (22, 24). The invention also proposes a device for implementing the method.
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
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29C 33/38 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor characterised by the material or the manufacturing process
B29C 33/48 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles with means for collapsing or disassembling
B29C 70/54 - Component parts, details or accessories; Auxiliary operations
B29C 70/76 - Moulding on edges or extremities of the preformed part
B29C 70/86 - Incorporating in coherent impregnated reinforcing layers
The invention concerns a connecting device for a conduit (201) in piping, including: - a conduit (201); - a connection (231) including the means (255) to carry out a sliding link, sealed longitudinally with an extremity of the conduit (201) according to a path d parallel to the axis (210), said to be longitudinal, of the said conduit (201); - the means (265) to stop the first extremity of the conduit (201) in translation in the connection (231) for a path d1 lower than d. The means (265) to stop in translation are retractable for defined compression conditions of the conduit (201).
F16L 27/12 - Adjustable joints; Joints allowing movement allowing substantial longitudinal adjustment or movement
B64D 33/00 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for
B64D 37/00 - Arrangements in connection with fuel supply for power plant
F16L 3/16 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets with special provision allowing movement of the pipe
F16L 37/50 - Couplings of the quick-acting type allowing movement of the parts joined
F16L 37/56 - Couplings of the quick-acting type for double-walled or multi-channel pipes
F16L 39/04 - Joints or fittings for double-walled or multi-channel pipes or pipe assemblies allowing adjustment or movement
17.
DETERMINATION PROCESS FOR AIRCRAFT AIR SPEED AND AIRCRAFT EQUIPPED WITH IMPLEMENTATION MEANS
The invention aims to provide a sufficiently precise alternative airspeed information for an aircraft, usable in the part of the flight area where the means based on the measurement of the incidence are not appropriate. To do so, the invention uses pressure information delivered by the aircraft engines usable in the systems in the plane level of the aircraft. In particular, the aircraft is equipped with engines (1) composed of at least one nacelle (30) and a compressor/turbine hitch (2, 3; 12, 14, 16, 18), the aircraft including at least one data treatment unit (40) and at least one system at plane level. At least one engine (1) of the aircraft is equipped with at least one ambient air static pressure sensor (P1) under the nacelle (30), static air pressure sensors (P2, P3) arranged at the engine level, one pressure sensor (P4) at the compressor output (12), at least one rotation speed sensor (N1, N2), for example a fan (10) and ambient air temperature sensors (T1) of the input air (E1) or at a compression point (T1). The treatment unit (40) is then capable of delivering the airspeed determination data at one least one of these systems at the plane level, based on the measurements carried out by these detectors at the engine level.
The invention concerns an aircraft including: - a drainage system for condensation formed in the aircraft between an upper zone of the aircraft where the condensation forms and a lower zone where the condensation is collected, - a thermal insulation mattress (16) with two opposite faces where a first face (16a) is located opposite the fuselage. The drainage system includes a drainage conduit (18, 24) which extends into the upper zone along the second opposing face (16b) of the insulating mattress and which is connected to a crossing device (30; 40) of the mattress placed under the upper zone. The drained condensation thus crosses the mattress using the crossing device and is then drained towards the lower zone, between the first face of the mattress and the fuselage.
Rear aerodynamic fairing of a linking strut ensuring an engine is fastened under the wing of an aircraft, comprising a lateral inner panel and a lateral outer panel, assembled to one another using a lower panel such that: - the lower panel includes, on either side of the lateral panels, an extension forming an aerothermal barrier intended to channel the hot air flow exiting the aircraft engine during use; - the inner lateral panel is assembled to the lower panel using a first angle iron secured in an angle formed by the lower panel and the lateral inner panel, - the outer lateral panel is assembled to the lower panel using a second angle iron secured in an angle formed by the lower panel and the outer lateral panel.
Assembly including two aircraft fuselage sections assembled end to end longitudinally and which each includes a skin as well as longitudinal stiffeners and at least one assembly device including at least one bracket which includes two bracket parts to attach two corresponding longitudinal stiffeners essentially aligned with one another and each respectively part of the said fuselage sections and at least one removable linking element. The two bracket parts each include at least one longitudinal footing, a core and a transverse head rest, the two bracket parts distanced from one another longitudinally by a circumferential frame element. Each of the two corresponding longitudinal stiffeners includes a respective footing and core.
- Device for the transverse retention of a longitudinal test specimen and apparatus for the mechanical characterization in bending of such a test specimen. - According to the invention, the device (1) for the transverse retention of a longitudinal test specimen (2), comprising a web (2A) and at least one main flange (2B) connected to one end of said web, comprises: - a body (3) in which there is formed a housing (4) which is shaped to house said test specimen (2) transversally and which has a main bearing face (5) against which one face of said web (2A) is intended to press; and - mobile clamping means (6) formed so as to press said web (2A) against said main bearing face (5).
A joint comprising first and second rib components, the first rib having an abutment surface and a plurality of lugs disposed adjacent to the abutment surface; and the second rib having an abutment surface and a plurality of lugs disposed adjacent to the abutment surface, wherein the abutment surfaces of the respective ribs are abutting and joined with a tension joint, and wherein the lugs are pinned to form a pinned lug joint by aligning respective holes in the plurality of lugs such that the lugs of the first and second ribs are interleaved and have one or more pins passing through the aligned lug holes. The joint may be used for attaching a wing tip device to the outboard end of an aircraft wing.
The invention relates to an aircraft with reduced environmental impact including two turboprop engines (10) mounted on the back of the aircraft at the rear thereof, and one acoustic masking device per turboprop engine. The acoustic masking device includes a masking element (32), such as a flap (33), which can move between a position in which it is retracted into the wing, and a position in which it is extended toward the rear of the wing. In said extended position, the flap (33) intercepts the noise area (18) generated toward the front by the turboprop engine, so as to reduce the perception of noise on the ground.
An aircraft includes many stiffened structures. In some cases, these stiffened structures must be crossed by a conduit essentially parallel to the plane of the stiffeners. This becomes problematic when the conduit has a section above the mesh of a structure and/or when it must be connected to the structure. The object of the invention is a stiffened structure including stiffeners and knots at the level of which three stiffeners of different directions cross. The faces of the stiffeners are arranged in planes perpendicular to a same plane. The stiffened structure includes at least one orifice and is characterized in that the said orifice is surrounded by a hexagonal belt made of peripheral stiffeners.
The invention relates to a fish joint device for securing first and second structural elements (1, 2) of an aircraft to each other, said device comprising two fish plates (20, 30) arranged either side of the structural elements (1, 2) and fixed thereto by means of a set of cross-members (11) for fixing fish plates. According to the invention, the fish joint device comprises at least one intermediate fish joint plate (40, 50) that is arranged between one of said fish plates (20, 30) and said structural elements (1, 2) partially covered by said intermediate plate, and is fixed thereto by a set of cross-members (11) for fixing intermediate plates, including the set of cross-members (11) for fixing fish plates.
This invention concerns a redundant signal (X1, ..., XN) processing system (12), an associated process, as well as an aircraft (2) including such a system, in order to monitor and passivate erratic or oscillating failures affecting the sources (20) of these redundant signals. The system includes: - a calculation module (120) for a current useful signal (U) based on the redundant signals; - a monitoring/passivation module (124) apt to detect an erroneous signal, and based on a criterion (T), to set aside the said erroneous signal from the calculation; - and as soon as an erroneous signal is detected, the means to toggle (122, 1220), to a freeze mode (M2), freezing the output useful signal and to return to an emission mode (Ml) as soon as no erroneous signal is detected, where the current useful signal is emitted as the useful output signal (X).
The aim of the invention is to allow the set of doors to be adjusted without increasing the number of landing gear retraction/extension manoeuvres, by incorporating a specific entrance in the fairing (2) of the landing gear bay (4) that provides access to the different accessories (14, 17a, 17b) used to adjust the doors (21-25). The method of the invention consists in: providing a removable bodywork element (5) which is inserted into the fairing (2) of a landing gear bay, surrounded by a stiffening structure (7) and of sufficient dimension such that the adjustment levers (14, 17a, 17b) can be reached from the outside; accessing the inside of the bay (4) by retracting the removable bodywork element (5); and using the levers (14, 17a, 17b) to adjust the different doors (21 to 25) in a flush aerodynamic position, as well as the clearances between said doors. In one embodiment of the invention, one door (25) comprises a central bodywork portion (5) and a peripheral portion (7) including stiffening means (70). The peripheral portion (7) is surrounded by a seal (30) that bears on the other doors (21-24) via the stiffening means (70) and the central portion (5) is coupled to the peripheral portion (7) by removable separation means (6).
The specific purpose of the invention is the synchronization of a local clock for apparatus in a communications network based on synchronization frames emitted by reference equipment and transmitted by a switch. After having received (700) a synchronization frame emitted by the reference equipment and including its emission date, the switch transmits (730) the said synchronization frame received after a fixed delay, at least on average, from the reception of the said synchronization frame by the said switch (725). The local clock of the equipment is synchronized after the reception of the said synchronization frame by the equipment according to the said emission date, the reception date of the said synchronization frame and the said delay.
The invention concerns the balancing of a simulation platform (1) including a non-linear aeronautical system (Mi, 20, 100). The balancing is based on: a servitude (200, 300) of an output value (Y ,Y (-1)) of the system over an instruction value (Y(1)k, cons, Y cons), including a first filter (300) creating the inversion of a simplified model (F m, G m) of the system to generate a said command (U) for the system from a correction information of the model (.DELTA.corr) estimating a modelling error (.DELTA.m) of the simplified model (F m, G m) and difference value (v) between the output value and the instruction value, and a corrector modelling loop including a second filter (400) implementing the said simplified model to calculate the correction information of the model to be provided for the first filter from the said control and the said output value.
The invention relates to a control member (A) comprising actuation members (F) such as a rudder bar and brakes for a pilot in an aircraft, said actuation members comprising a rotary shaft (2) that is rotatable around an axis (40). Said shaft is rotated during an action of the pilot on the actuation members (F), in particular the brake pedals, by means of mechanical transmission between said shaft and said actuation members. Said control member (A) also comprises a device (1) for generating resistive torque, which is suitable for being attached along the rotary axis of the control member, the device for generating resistive torque comprising a friction disk (3, 4, 5) that is axially mounted and secured to the rotary shaft (2). A first annular track (6, 7, 8) is attached to a frame element (9) of the control member and to bearing means (10, 11, 12).
B64C 13/46 - Transmitting means with power amplification using fluid pressure with artificial feel
G05G 5/03 - Means for enhancing the operator's awareness of the arrival of the controlling member at a command or datum position; Providing feel, e.g. means for creating a counterforce
31.
METHOD, COMPUTER PROGRAM AND DEVICE FOR SUPERVISING A SCHEDULER FOR THE MANAGEMENT OF PROCESSING TIME SHARING IN A MULTITASK DATA-PROCESSING SYSTEM
The invention in particular has as an object supervising a scheduler for the management of processing time sharing in a multitask data-processing system comprising a computation unit having a standard execution mode and a preferred execution mode for executing a plurality of applications. The execution time for the said plurality of applications is divided into a plurality of periods and a minimal time for access per period to the said computation unit is determined for at least one application of the said plurality of applications. For at least one period, the said preferred execution mode is associated with the said at least one application and the said at least one application is executed according to at least the said minimal time for access to the said computation unit. For the said at least one period, the said standard execution mode is associated with the applications of the said plurality of applications and at least any one of the applications of the said plurality of applications is executed.
Equipment for manufacturing a composite panel, particularly for aircraft fuselage. The equipment (1) includes a mold (2) featuring a fixed part (4) and a plurality of removable modules (6) capable of being arranged on this fixed part (4).
AIRBUS OPERATIONS (SOCIETE PAR ACTIONS SIMPLIFIEES) (France)
Inventor
Blot, Philippe
Hochart, Benoit
Bechtel, Stephane
Abstract
The invention relates to a method for manufacturing an aircraft part, comprising: arranging, in a stack including a preform (14), at least one layer (22, 29) of a material that has a perviousness to a predetermined resin that is less than that of a portion of the perform that is closest to the layer; and arranging, on one side of the layer (22, 29) opposite the perform, a sensor (24) for the presence of the resin.
The invention relates to an aircraft turbine engine propeller, comprising a plurality of blades (1) mounted on a hub (2) that is rotatable about a rotational axis (I-I) of the propeller, the aim of the invention being to reduce the risks of losing the blade after impact from a foreign body. To accomplish the aim, the blade includes a recess (30) leading radially inside at the root (20) thereof. A blade-retaining device (50) includes movable retaining means (53A, 53B) located in said recess (30) and movable from a retracted position to a deployed position for retaining the blade, and vice versa.
The invention relates to a device for protecting piping from lightning, characterized in that the piping consists of tubular metal sections (1, 2, 3) connected to each other by connection parts (6a, 6b, 4a, 4b, 51, 52), a first portion (4a, 4b, 51) of the connection parts being plastic material parts and a second portion (6a, 6b, 52) of the connection parts being metal parts.
F16L 25/02 - Construction or details of pipe joints not provided for in, or of interest apart from, groups specially adapted for electrically insulating the two pipe ends of the joint from each other
H05F 1/00 - Preventing the formation of electrostatic charges
36.
TAPPED INSERT FOR CYLINDRICAL BORE AND INSTALLATION EQUIPMENT FOR SUCH INSERT
The present invention relates to a tapped insert for cylindrical plain bore (30) of a part A1 comprising a cylindrical body (10), characterized in that said cylindrical body (10) comprises an inner bore (11) of frusto-conical shape, and at least one longitudinal slot (12, 12') extending into the cylindrical body (10) and communicating with said frusto-conical inner bore (11) and characterized in that the tapped insert also comprises a core (20) of frusto-conical shape complementary with said frusto-conical inner bore (11) of the cylindrical body (10), said core (20) comprising a inner tapped aperture (21), said cylindrical body (10) and said core (20) being such that said tapped insert can be holded in position and locked in said cylindrical bore (30) to equip said cylindrical bore with an inner taping.
F16B 13/06 - Dowels or other devices fastened in walls or the like by inserting them in holes made therein for that purpose with parts gripping in the hole or behind the reverse side of the wall after inserting from the front combined with expanding sleeve
F16B 2/04 - Clamps, i.e. with gripping action effected by positive means other than the inherent resistance to deformation of the material of the fastening internal, i.e. with spreading action
37.
METHOD FOR MAKING A PROFILE SECTION MADE OF COMPOSITE MATERIAL FIBERS AND HAVING A CROSS-SECTION IN THE SHAPE OF A HALF OF A NAIL HEAD OR OF A NAIL HEAD
The invention relates to a method for making a profile section made of fibers (32) and intended for filling a space, having a cross-section in the shape of a nail head or of a half of a nail head resulting from the assembly of preforms in a T shape, at least one of said performs having a radius of curvature at the junction area of said preforms, characterized in that the method consists of feeding fibers (28) coated with a powder to a preforming station (30) at which said powder is activated and said fibers (28) move in a passage, the shapes of which are adapted to those of the desired profile section, and then moving said preformed fibers in a shaping station (34) at which said fiber profile section (32) is cooled and shapeed by passing into a cavity, the shapes of which are adapted to those of the desired profile section.
B29C 70/20 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in a single direction, e.g. roving or other parallel fibres
B29C 70/52 - Pultrusion, i.e. forming and compressing by continuously pulling through a die
38.
METHOD AND DEVICE FOR INTEGRATING A RF GPS EMULATOR IN A SIMULATION ENVIRONMENT
The invention in particular has as an object a method and a device for integration of an RF GPS emulator in a simulation environment comprising a computer adapted for using a simulation timer and for receiving a signal of x_PPS type. For each pulse detected on the said x_PPS signal, a counter for number of detected pulses is updated (425) and the value of the said simulation timer is stored (425). The evaluation (415) of a given moment, in the clock of the said emulator, is performed according to the value of the said counter for number of pulses, the value of the said simulation timer and the said stored value of the said simulation timer. A step of verifying (430) the detected pulses preferably is implemented.
The invention relates to a system for selecting a datum representative of an air parameter, including a plurality of sources each capable of delivering a datum representative of a measured value of the parameter. The system includes: means for determining, for at least some sources, the number of other sources from the plurality in agreement with the source in question; means for selecting the datum selected from the data delivered by the sources for which the determined agreement number is the maximum. The invention also relates to an engine control system, to an aircraft including such systems, and to an associated method.
Le dispositif pour la détermination de la position d'une manette des gaz (2) selon l'invention comporte des capteurs de position (P, R) relies chacun a un calculateur de commande de vol primaire des capteurs de position relies chacun a un calculateur de contrôle moteur, lesdits capteurs de position étant répartis en au moins trois groupes de capteurs n'ayant pas de panne simple commune, et au moins un calculateur (PRIM), dit calculateur d'interface, présentant au moins une entrée pour recevoir des informations de mesure émises par les calculateurs reliés aux capteurs de position ainsi que des sorties vers le calculateur de contrôle moteur (EEC).
The data frequency analysis method comprises: a step (310) for inputting signals coming from a first sensor; a step (315) for inputting signals coming from at least a second sensor, each second sensor being positioned close to the first sensor so that the signals coming from each second sensor are strongly correlated with the signals coming from the first sensor; a step of estimating, for each sensor, a transfer function or model established from the combination of the signals from the first sensor and from each second sensor; and a step (320) of extracting the structural properties of the system from each of the estimated models.
ESTIMATION OF A CRITERION OF LOAD TO WHICH A STRUCTURAL COMPONENT OF AN AIRCRAFT IS SUBJECTED, AND ASSISTANCE FOR THE DETECTION OF A SO-CALLED "HARD" LANDING BY VIRTUE OF SUCH A CRITERION
The present invention relates to methods and systems for estimating a loading criterion relating to the load (C) experienced by a structural component (100, 102, 104) of an aircraft (10), and assistance with detecting a so-called "hard" landing. These methods involve measuring (50) parameters of said aircraft and calculating (54) at least one loading criterion for the loading of said structural component using at least one neural network (36) receiving said parameters as input. Assistance with detecting a hard landing then requires the determining (72, 76) of a time (t0, t-i) of impact (14, 16) of said aircraft on a landing strip (12) from said measured parameters, then estimating (74, 78) a plurality of said parameters at said determined time of impact so as to calculate (54) the at least one loading criterion relating to the loading of the structural component.
The invention relates to a device for reducing turboprop infrared emissions. According to the invention, fresh air (E) at the turboprop outlet is captured and mixed with the warm airflow (F) created by said turboprop to lower the temperature.
The invention relates to a method for assembling a piece of metal material (1) to a piece of composite material (2), with the assembling of the said pieces taking place by means of an attachment member (11), characterized in that the method comprises the following steps in which: - the two pieces are pre-drilled so as to form a gap (6) of diameter D1 in the metal piece (1) and in the piece of composite material (2), - an insert is applied (9) that is of variable diameter in the direction of the thickness of the pieces in said gap (6), said insert comprising a main hollow body (12) and at least one loop (10) intending to at least partially support against the outside surface (101) of the piece, said insert (9) intending to accommodate the attachment device (11), - residual compressive stresses are generated by radial expansion on the inner wall of the insert, by means of an expansion tool (7) that has an outside diameter engaging with the inserts inner diameter; these residual compressive stresses are generated in at least one area of the two pieces along the circumference of the insert (9), - the gap in the main body (12) of the insert is created such that the insert has a final inside diameter Dfinal that is substantially larger than the outer diameter of the attachment member (11) so as to allow some play between the inserts inside wall and the outside surface of the attachment member, and the attachment member (11) is positioned inside the insert to keep the metal and composite pieces together.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
INSTITUT NATIONAL POLYTECHNIQUE DE TOULOUSE (France)
Inventor
Iturriz , Marcelo Fernando
Richardeau, Frederic
Meynard, Thierry
Helali, Hassan
Abstract
The invention relates to a power rectifying circuit (30) for an electric current signal supplied by an alternating power source (12), which comprises: two separate switching assemblies (30a, 30b) adapted to be connected to a power terminal (31) of the source, wherein at least one switching assembly includes a plurality of boost cells (310a, 310b, 320a, 320b) in cascade, each boost cell including a diode (312a, 312b, 322a, 322b), a switch means (314a, 314b, 324a, 324b) and a capacitor (316a, 316b, 326a, 326b), the so-called terminal capacitors (326a, 326b) of the two terminal boost cells (320a, 320b) of the switching assemblies having one terminal in common. The circuit may particularly include two assemblies of boost cells. The invention can be used in electric systems onboard aircrafts.
H02M 7/10 - Conversion of ac power input into dc power output without possibility of reversal by static converters using discharge tubes without control electrode or semiconductor devices without control electrode arranged for operation in series, e.g. for multiplication of voltage
H02M 7/25 - Conversion of ac power input into dc power output without possibility of reversal by static converters using discharge tubes with control electrode or semiconductor devices with control electrode using devices of a triode or transistor type requiring continuous application of a control signal using semiconductor devices only arranged for operation in series, e.g. for multiplication of voltage
46.
AIRCRAFT NACELLE INCLUDING HOT AIR DISCHARGE MEANS
The invention relates to an aircraft nacelle that comprises a discharge means (50) for establishing a communication between the inside of the nacelle and the outside thereof, and a junction area between two members (54, 56), and at which the two members are superimposed, both members each having a surface defining the outer surface of the nacelle on either side of the junction area, characterised in that said discharge means (50) includes a dual-portion shim (52), i.e. a first portion provided between the superimposed panels and including at least one protruding and/or recessed member for forming a passage between said panels in order to establish a communication between the inside and the outside of the nacelle, and a second portion at the level of only one panel, i.e. that located inside at the junction area, in order to protect the same against excessive temperatures.
The invention relates to an aircraft leading edge including, as an extension, an aerodynamic surface at which flows an aerodynamic flow and where are provided air outlets (30) for preventing the separation of said aerodynamic flow, the air outlets (30) being arranged into at least two rows substantially parallel to the leading edge and in an offset manner for at least two successive layers, characterised in that it comprises at least one shim (36) provided between two walls defining the aerodynamic surface, said shim having, on the one hand, an outer surface (38) arranged as an extension of the aerodynamic surface (26), a first inclined surface (40) in contact with the first wall defining the aerodynamic surface, and a second inclined surface (42) in contact with the second wall defining the aerodynamic surface and, on the other hand, protruding and/or recessed patterns formed at the inclined surfaces (40, 42) and alternating from one face to the other, that allow the passage of air on either side of the aerodynamic surface.
The invention relates to a method for controlling the thermal effluents of an aircraft including an airframe (110) and at least one propulsion unit (112), wherein said at least one propulsion unit (112) includes a turbomachine (116) supplied with fuel via a fuel supply circuit (122) extending from a tank (124) provided at the airframe (110), said airframe (110) including at least one thermal effluent source (134), characterised in that it comprises scattering at least partially, at the level of at least one propulsion unit (112), the thermal effluents generated at the airframe (110) by using the fuel used for supplying the turbomachine (116) as a heat-carrier fluid.
According to an exemplary embodiment of the present invention, a method for mounting a wing of an aircraft to a fuselage of the aircraft is provided, in which a difference between a vertical target position and a vertical actual position of a mounting point is determined. Then, on the basis of the determined difference, a readjustment of the wing is performed.
The invention relates to a system for transmitting stresses at the intersection between a reinforcing frame and a stiffener in a fuselage and to a method for producing one such system. An aircraft fuselage includes stiffeners (4) interrupted by reinforcing frames (5). At said interruptions, the stresses experienced by the stiffeners (4) must be transmitted despite the presence of the reinforcing frames (5). Known solutions are difficult to manufacture and install and do not provide a simple stress transmission path when produced solely from composite materials. In order to solve this problem, the invention includes the creation of battens (15) that cover the ends (16) of consecutive sections (6) of a stiffener (4) and the skin (7) of the fuselage (1), extending under the reinforcing frame (5) in order to join said two consecutive sections (6).
The invention mainly relates to a device for attaching an aircraft engine that comprises a rigid structure (10) and a means for attaching the engine on said rigid structure, said attachment means including a rear engine fastener (8) and a device (9) for collecting the thrust forces generated by the engine, the rear engine fastener (8) being attached on the rigid structure by two side brackets attached to the rigid structure, said force collecting device (9) including two connecting rods (26a, 26b) mechanically connected to a rudder bar (28) by a mechanical link, a connection bracket (49) attached on the rear engine fastener (8) and mechanically connected to the rigid structure (10) by a thrust axis (60) wherein the side brackets include abutment means limiting the tilting of the rudder bar (28) in the case of breakage of a connecting rod (26a, 26b) and ensuring the transmission of the thrust forces to the rigid structure (10).
The invention relates to methods and devices for the non-intrusive identification and communication between a server node (259 and at least one client node (10) in a communication network (5). An identification message is first transmitted to at least one client node, the identification message including at the application level a duplicate of the addressing data (the physical address and the internet address of said server node) also contained in the MAC and IP layers. The reception of at least one couple of addresses from at least one client node, wherein a couple of addresses from a client node includes a physical address and an internet address (according to the same duplication strategy), can be used for identifying the client node by the server. Similarly, the invention can be used for identifying in a client node at least one server node in the network.
H04L 61/103 - Mapping addresses of different types across network layers, e.g. resolution of network layer into physical layer addresses or address resolution protocol [ARP]
H04L 67/125 - Protocols specially adapted for proprietary or special-purpose networking environments, e.g. medical networks, sensor networks, networks in vehicles or remote metering networks involving control of end-device applications over a network
H04L 12/18 - Arrangements for providing special services to substations for broadcast or conference
The invention relates to a seat for a vehicle, that comprises a headrest (1) with a headstall (2) for supporting the head (51) of a passenger (5) in the seat, wherein said occupant may or may not wear a protection helmet (52) depending on the circumstances, characterised in that the headstall (2) can assume at least three stable positions: a first position for the cervical support of the head of the seat passenger in which the headrest (1) provides a cervical support function with or without wearing a helmet (52); a second position for the occipital support of the head of the seat passenger in which the headrest (1) provides an occipital support function without wearing a helmet (52); a third position for the occipital support of the head of the seat passenger in which the headrest (1) provides an occipital support function while wearing a helmet (52).
The invention relates to a method and a device for managing communication channels for data exchange from an aircraft (105) comprising means (115) for establishing at least one connection path between the aircraft and at least one ground entity (120) with a plurality of communication channels of at least two different types. After detecting an event such as the modification of a communication channel condition or a modification of the aircraft position, a communication configuration enabling the aircraft to receive or transmit data from or to the ground entity via at least one of said communication channels is determined. The means for establishing at least one connection path are then adapted based on the communication configuration. Advantageously, the transmission of at least one datum comprises determining the priority level of said datum and determining the type of communication channel of the connection path, the datum being transmitted if the determined type of communication channel is compatible with the priority level.
The invention relates to an aircraft comprising a cabin, on at least one wall of which are attached the first portions of several multisystem connection devices, each device comprising a male portion or plug (80) and a female portion or socket (81) capable of fitting into each other, wherein each male and female portion comprises: - first means for connecting to a liquid or gas supply of the aircraft, - second means for connecting to a supply of electricity and/or electric signals of the aircraft.
B64D 13/00 - Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space
F16L 25/01 - Construction or details of pipe joints not provided for in, or of interest apart from, groups specially adapted for realising electrical conduction between the two pipe ends of the joint or between parts thereof
H01R 13/00 - ELECTRICALLY-CONDUCTIVE CONNECTIONS; STRUCTURAL ASSOCIATIONS OF A PLURALITY OF MUTUALLY-INSULATED ELECTRICAL CONNECTING ELEMENTS; COUPLING DEVICES; CURRENT COLLECTORS - Details of coupling devices of the kinds covered by groups or
56.
METHOD FOR PRODUCING AN ACOUSTICALLY RESISTIVE STRUCTURE, RESULTING ACOUSTICALLY RESISTIVE STRUCTURE AND SKIN USING ONE SUCH STRUCTURE
The invention relates to a method for producing an acoustically resistive structure which can be mounted on a honeycomb structure such as to produce a skin for acoustical treatment, said acoustically resistive structure including at least one porous layer (34) and at least one reinforcing structure (36) which are glued together. The invention is characterised in that the method comprises the following steps in which: an amorphous adhesive is applied to the at least one reinforcing structure (36), the reinforcing structure (36) is subsequently perforated or micro-perforated, and the at least one porous layer is applied against the surface of the reinforcing structure (36) that was coated with amorphous adhesive.
The invention relates to a device for measuring the flow speed of a fluid as well as its direction and its orientation, based on a thermal-sensor measurement principle, that comprises at least three flow measurement probes (1, 1a,..., 1f) each having a sensitive member (2) and an obstacle (3) masking a predetermined measurement area of the sensor, characterised in that the flow measurement probes are attached on carrier posts (7), said carrier posts (7) defining said obstacles (2) that form masking members (4) for an angular sector of the flow measurement probes facing the sensitive member of said probes.
G01P 5/10 - Measuring speed of fluids, e.g. of air stream; Measuring speed of bodies relative to fluids, e.g. of ship, of aircraft by measuring thermal variables
G01P 5/12 - Measuring speed of fluids, e.g. of air stream; Measuring speed of bodies relative to fluids, e.g. of ship, of aircraft by measuring thermal variables using variation of resistance of a heated conductor
G01P 13/00 - Indicating or recording presence or absence of movement; Indicating or recording of direction of movement
G01P 13/04 - Indicating positive or negative direction of a linear movement or clockwise or anti-clockwise direction of a rotational movement
58.
COATING FOR ACOUSTIC TREATMENT HAVING A HOT AIR ICING PROCESSING FUNCTION
The invention relates to a coating for the acoustic treatment of the surface of an aircraft, especially at the level of a leading edge such as the air intake of an aircraft nacelle, said coating including an acoustically resistive layer, at least one cellular structure (30) and a reflective layer, said cellular layer (30) including a plurality of ducts giving, on the one hand, onto a first imaginary surface (34) and, on the other hand, onto a second imaginary surface (36), characterised in that the cellular structure (30) includes cut-outs or openings (38) formed at the side walls of some ducts for establishing a communication between adjacent ducts in order to create a network of communicating ducts that isolate at least one non communicating duct or group of ducts, one at least of said communicating ducts being connected to at least one hot-air inlet (40).
The invention relates to a nacelle, for an aircraft jet engine (10) having a high bypass ratio, in which a jet engine having a longitudinal axis (X) is mounted, the nacelle (12) including a wall (24) concentrically and at least partially surrounding the jet engine and defining with the latter an annular duct (26) for a fluid internal flow having, at the downstream end of the nacelle wall, a flow outlet passage section, characterised in that the nacelle includes displacement means (40) for controllably displacing a portion (24b) of the nacelle wall in order to modify the section of the flow outlet passage, wherein said displacement generates in the nacelle wall at least one longitudinally extending opening (28), the nacelle including a device (30) for forming a fluid barrier (fi) extending along a portion at least of the longitudinal extent of said at least one opening (28) in order to counteract the natural exhaust through said at least one opening of a portion of the so-called trailing flow.
F02K 1/28 - Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto using fluid jets to influence the jet flow
B64D 33/04 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of exhaust outlets or jet pipes
F02K 1/72 - Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
60.
PROCESS FOR MANUFACTURING A PANEL MADE OF A THERMOPLASTIC COMPOSITE
The present invention relates to a process for manufacturing a panel made of a composite with the aid of tooling having a support (1) on which a lay-up step is carried out, said step consisting in producing a stack (18) of plies of fibre prepregs (20) followed by a step of consolidating the stack with the aim of obtaining the panel by means of a compacting plate (36) placed on top of this stack. According to the invention, the process is implemented in such a way that a first set (26) of polyimide films (28) partly overlapping one another is placed in contact with the stack of plies (18), between said stack and the compacting plate.
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
The invention concerns a device for detecting and eliminating a layer of ice formed on the surface of an aircraft structure (6) or a liquid that has infiltrated inside of a structure and/or into the material of the structure. According to the invention, the device comprises at least one pair of subnetworks of conductive elements (2a, 2b), each subnetwork comprising at least one series of conductive elements (3), said subnetworks being arranged in such a way that the interfitting between the conductive elements (3) of the first subnetwork and the conductive elements (3) of the second subnetwork forms a network of capacitive sensors (1), said conductive elements being embedded in an insulating material (4), each subnetwork of conductive elements being incorporated into a flexible substrate (5, 5a, 5b) so that the entire assembly forms a flexible covering.
The invention concerns a conduit incorporated into the tubing of an aircraft characterized in that it is made from a composite material comprising a fiber-reinforced geopolymer resin matrix.
F16L 9/14 - Compound tubes, i.e. made of materials not wholly covered by any one of the preceding groups
C04B 28/00 - Compositions of mortars, concrete or artificial stone, containing inorganic binders or the reaction product of an inorganic and an organic binder, e.g. polycarboxylate cements
F16L 9/18 - Double-walled pipes; Multi-channel pipes or pipe assemblies
According to the invention, the air intake (2) is made of a piece (23) of fibre/resin composite material, reinforced on the side of the flange (24) by a reinforcing ring (26).
According to the invention, high lift leading-edge slats (13) and high lift flaps (14) are simultaneously deployed, when the speed of the aircraft is equal to and less than an AES threshold.
The purpose of the invention is a safety device for a semiconductor switch controlling a secondary line (2) provided with a fuse (3) on the input side of the switch device that comprises component (4) to create a short circuit in the secondary line, so as to cut off the fuse (3), and means (5, 6, 7) of activating this component (4) following detection of an overcurrent in the secondary line.
A tool (10) for cold expansion of holes comprises a body (14) which has at least a cylindrically shaped part (16) having an inner hole (24) in axial direction of the body (14). The body (14) comprises along the circumferential direction a plurality of members (18, 18a, 18b, 18c) which are movable in radial direction of the body (14), such that the outer diameter of the cylindrically shaped part (16) of the body (14) can be changed. Furthermore a core (26, 26c) is provided which is insertable into the inner hole (24), wherein at least one of the core (26, 26c) and the inner hole (24) comprises a portion (28) of changing diameter and the inner hole (24) and the core (26, 26c) get into contact with each other upon insertion of the core, such that in dependency of the depth of the insertion of the core (26, 26c) into the inner hole (24) in axial direction the members (18, 18a, 18b, 18c) move radially outwardly and the outer diameter of the cylindrically shaped part (16) of the body (14) increases.
The invention relates to an aircraft wing arrangement comprising a wing and at least one strut for attaching the aircraft engine suspended under the wing and comprising a front zone situated projecting towards the front with respect to a leading edge of the wing. According to the invention, the front zone comprises a lateral projection delimiting a lateral airflow channel while being projected towards a proximal end of the wing. One technical problem addressed by the present invention relates to aerodynamic continuity and efficiency for wings having an aircraft engine closely suspended under the wing.
According to the invention, a sealed articulation device (27) is mounted on at least one (14) of the branches (14; 18, 9) of said hot-air circulation circuit in the form of a bracket so as to provide said bracket with capabilities of deforming about said articulation device (27).
The device (1) comprises a means (3) for determining the current weight of the aircraft, a means (4) for calculating a reduced thrust value which is proportional to the said current weight, a means (5) for determining at least one engine command, so that the engines deliver a thrust equal to the said reduced thrust value, and a means (7) for applying this command to those engines of the aircraft which have not failed.
The invention concerns a system (1) comprising a steering vane (2) which can turn within a rudder deflection range which is limited by first and second rudder deflection limits, and means (12) for asymmetrically varying said first and second rudder deflection limits based on the current values of the aircraft flight parameters.
The subject of the invention is a device for protecting and guiding at least one cable (22) connecting an openable window (10) for closing an opening of an aircraft cabin to its mounting (14), and is characterized in that it comprises a cable-supporting chain (26) which is deformable within a winding plane and is able to protect and guide said cable (22), and on either side of the cable-supporting chain (26) a first mounting (28) hinged to the openable window (10) in such a way as to be able to pivot about a hinge axis (36) extending approximately parallel to the translational movement of disengagement and a second mounting (30) hinged to the mounting (14) in such a way as to be able to pivot about a hinge axis (38) extending approximately parallel to the translation movement of disengagement, and in that the first and second mountings (28, 30) have an L-shaped cross section.
F16L 3/015 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets for supporting or guiding the pipes, cables or protective tubing, between relatively movable points, e.g. movable channels using articulated- or supple-guiding elements
According to the invention, on at least one of the faces (1E, 1I) of said leading edge (1), forward of the partition (3), an action is exerted on the fluid flow (EF) so as to cause the latter to switch from the laminar state (EFL) to the turbulent state (EFT) in order to increase the thermal exchanges and minimize the thermal stresses on said leading edge.
The invention relates to an aircraft (10) comprising: - a fuselage (12), - two wings (14, 16) to which engine nacelles are attached and which are each connected laterally to the fuselage, one on each side thereof, by a central fairing, the central fairing (18, 20) comprising, facing each wing, two opposed surfaces connected one to the suction face side and the other to the pressure face side of the wing in question and which extend longitudinally along the fuselage, characterized in that at least one of the two surfaces has at least one local geometric deformation (50; 52; 54, 56, 58) which is designed to generate lateral aerodynamic disturbances on the central fairing toward the wing with a view to controlling the flow of air over the wing.
An engine speed control device comprises an engine speed control lever (2) generating an engine speed instruction (POS) and means for delivering the indicative information (.theta.e) on the actual engine speed. Actuating means (8, 10) are designed in such a way that a mechanical force can be applied on the lever (2) according to the difference between the instruction (POS) and the indicative information (.theta.e) on the actual engine speed. Said device is suitable for an aircraft. The engine speed control method is also disclosed.
The invention relates to a mount for aircraft turbojet engine, comprising a central box (22) formed by the assembly of two side panels joined via transverse ribs (23), the mount also comprising two side boxes (24a) secured to a forward part of the box (22) and arranged either side thereof, each side box having an aft closure frame (46a). According to the invention, one of the ribs (23) forming the box (22) is equipped with two side extensions (52a) made in a single piece with this rib (23) and respectively projecting from the two side panels outwardly from the box (22), the two extensions (52a) being fixedly mounted on the frame (46a) of each of the two side boxes, respectively.
The invention concerns a traction pad to be bonded for coating adhesion or surface cohesion tests by tearing comprising a counterbore designed to control the thickness of the adhesive film. Support elements such as wedges or a border maintain the base of the counterbore at the required distance from the tested surface and openings are provided for evacuating the adhesive prior to its controlled hardening when the traction pad is applied on the tested surface.
The invention relates to an aircraft engine assembly comprising a turbojet engine, an engine mount and a plurality of engine attachments. The latter comprises a first and second forward attachment (6a) fixed to the fan case and located symmetrically in relation to a plane defined by a longitudinal axis of the turbojet engine and a vertical direction thereof, both forward attachments being designed so as to transfer loads exerted in a longitudinal direction of the turbojet engine and in the vertical direction thereof. Moreover, each of the two forward attachments comprises a shear pin (36) mounted on the case (12) and going through the two walls (32a, 32b) of a clevis (30) fixed on the engine mount.
This invention relates to a device for suspension of an engine, designed to be inserted between an aircraft wing and this engine, the device comprising a rigid structure (8) and a forward aerodynamic structure (24) mounted fixed on said structure (8) through suspension means, the forward aerodynamic structure being designed to carry engine fan casings and to be inserted between the rigid structure and the wing. According to the invention, the suspension means comprise at least one adjustable length connecting rod (32) of which one end is mounted on the rigid structure (8), and the other end is mounted on the forward aerodynamic structure (24).
The invention relates to a novel rear attachment (10) for an aircraft engine assembly. The inventive attachment (10) comprises a double- boomerang structure provided with two three-point shackles (12, 12') which are doubled in such a way that only four attachment points (22, 24, 22', 24') are provided on a mast (6) and two attachment points are provided on the engine (3), wherein the Fail Safe function is provided by a dual structure (12a, 12b). The inventive concept does not sacrifice security criteria by
The invention relates to an aircraft engine assembly (1) comprising an engine (2) and an engine mount (4), this engine mount including firstly a rigid structure (10) comprising a box (24) and secondly a mounting system (11) positioned between the engine and the rigid structure, this mounting system notably comprising an aft attachment (8), the assembly also being provided with a thermal protection system (58) for the box, comprising a preferably ventilated duct (60) forming a heat barrier and extending between this box and the engine. According to the invention, the duct extends rearward beyond the aft engine attachment.
The invention concerns an engine assembly (1) for an aircraft comprising an engine and an engine mounting structure (4) the engine, the assembly also comprising a heat exchanging system (104) provided with a second outlet (122a) located between the structure box and the engine, rearwards relative to the rear engine attachment (8). Moreover, the second outlet (112a) of the heat exchanging system is arranged on a second outlet duct (124) connected to the exchanger (114) and passing through a structural unit (34) mounted fixedly on a structure box (24), between the latter and the engine (2).
The invention concerns a turbofan provided with a pre-cooler. The invention is characterized in that to evacuate the heated cool air stream (24), at least one discharge pipe (25) is arranged in the chamber (12) and connects the pre-cooler (18) to at least one discharge orifice (26) provided in the inner fairing (10), in output of the exhaust nozzle (3) and at least more or less opposite the wing (16).
B64D 13/08 - Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space the air being conditioned the air being heated or cooled
F01D 9/06 - Fluid supply conduits to nozzles or the like
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
83.
METHOD TO MOUNT AN AIRCRAFT ENGINE ON A RIGID STRUCTURE OF AN ENGINE MOUNT
The invention concerns a method to mount an aircraft engine on a rigid structure of an engine mount, comprising a securing step, on the structure, to secure an attachment body (46) for an engine attachment, this attachment body being intended to lie against a contact surface of the rigid structure oriented forwardly, and secured to this structure by at least one shear pin (68) passing through a first and a second primary orifice, the securing step of the body (46) comprising the following operations: - pre-positioning the engine relative to the engine mount for the purpose of placing the first orifice opposite the second orifice; and - inserting a pin fitted with a convex centering head through the two orifices, so that this pin forms the shear pin.
The invention concerns an engine assembly (1) for an aircraft comprising an engine (6) and a device for locking (4) the engine, the locking device being designed to be interposed between a wing system of the aircraft and the engine, and comprising a rigid structure (8) as well as means for locking the engine (6) on said rigid structure, the locking means consisting of a first fastener (10) as well as a second fastener (12). The invention is characterized in that the first fastener comprises two lateral half- fasteners (10a) each fixed on a fan casing (18) of the engine, and the second fastener is fixed on a front part of a central casing (22) of said engine.
The invention concerns a strut for locking (4) the engine of an aircraft, the strut comprising a rigid structure (10) forming a structure box closed at the front by a front closing rib (36), as well as a system for mounting the engine including in particular a front engine fastener (6) comprising a fastener body (46) fixedly mounted on the rigid structure (10). The invention is characterized in that the fastener body (46) is pressed against a front surface (38), and the mast (4) comprises at least one shear pin (68) passing at least partly through the fastener body (46) as well as the front closing rib (36).
The invention concerns a perforating device comprising a cutting tool (10), means for rotating (11) the cutting tool (10) and means for translating (12, 14) the cutting tool (10), the ratio between the rotational speed and the translational speed being variable during rotation of the cutting tool (10). It comprises a gear train (16, 17; 18, 19) for synchronizing the rotating means (11) with the translating means (12, 14). The speed ratio between a driving pinion (16; 18) of the gear train, secured in rotation to the rotating means (11) of the cutting tool (10) and a transmission pinion (17; 19) of the translating means is reversed at least once during one rotation of the cutting tool (10). The invention is useful for fragmenting the formed shavings.
The invention relates to a double crimping tool for fastening an electrical conductor with insulation to a contact element. To this effect the following are provided: a front crimping unit with a front drive cam plate and front crimping stamps for forming a front press section in the region of the stripped end of the conductor; and a rear crimping unit with a rear drive cam plate and rear crimping stamps for forming a rear press section in the region of the insulation of the conductor. The crimping stamps are displaced by rotation of the drive cam plates. By levers the two drive cam plates are connected to a shared force introduction element such that during the pressing procedure at first optionally the front or the rear drive cam plate is rotated by a predefined initial pressing angle while the other drive cam plate is made to rotate only after the first drive cam plate has attained the initial pressing angle.
The invention relates to a system (33) for connecting a crossbar to the rigid structure of an aircraft engine pylon, comprising an axle system (32) which is mounted to slide in a first passage (38) in a fitting (34) such that it can move along a first longitudinal axis in a first direction from a normal extracted position to a retracted position in which it is retracted into the fitting, and vice versa. The inventive system also comprises an axle extension member (60) which is supported by the axle system (32) and which can be moved in parallel with the longitudinal axis in a second direction from a normal retracted position to an extracted position in which it is solidly connected to the system (32) and projects out in relation to same, and vice versa.
The invention concerns an aircraft with electric flight control system, wherein a device enables the load applied on the wing system to be lightened during a lateral roll maneuver, by comparing the roll control (X) with a threshold (X%) and by filtering said control. Through such an action on the roll control, it is possible to anticipate the load increases occurring on the wing system, and to reduce same. The loads can thus be maintained below a maximum value of the wing system dimensioning, in case of too great roll control relative to the use provided for the plane, and therefore to limit oversizing of the wing system.
The invention concerns a sol for sol-gel coating of a surface, said sol comprising in weight percentage: a) 3 to 30 %, preferably 5 to 20 %, more preferably 7 to 15 %, more preferably still 8 to 15 %, and more preferably still 10 to 13 %, for example 10.8 % or 12 % of at least one zirconium, aluminum, or titanium organometallic compound; b) 5 to 50 %, preferably 5 to 40 %, more preferably 10 to 40 %, more preferably still 15 or 20 to 30 %, for example 22 % or 23 % of at least one organosilane compound; c) 1 to 15 %, preferably 2 to 10 %, more preferably 2 to 8 %, for example 5 % of at least one compound selected among acids, bases, glycols and ethoxyethanol; d) the rest to make up 100 % being demineralized or distilled water; the total content in a) and b) being higher than 30 %, preferably higher than 31.2, 31.5, 32 or 33 %, more preferably higher than 35 %, more preferably still higher than 40 %, more preferably still more than 50 %. The invention also concerns a kit comprising: a first container holding a first part A in liquid form including the constituents a), c) and d) of the sol; a second container holding a second part B in liquid form including the constituent b) of the sol. The invention further concerns a method for preparing a sol-gel coating using said sol and the resulting coating, and a substrate coated with at least one such sol-gel. The invention also concerns a method for preparing a coating comprising several layers on a surface of a substrate, at least one of said layers being a sol-gel coating prepared by the inventive method.
C23C 18/12 - Chemical coating by decomposition of either liquid compounds or solutions of the coating forming compounds, without leaving reaction products of surface material in the coating; Contact plating by thermal decomposition characterised by the deposition of inorganic material other than metallic material
The invention relates to an engine mount for an aircraft, to be placed between an engine and an engine mounting structure, said mount comprising a shackle (20) which is connected to a first fixing device (34) comprising a fork (36) provided with two arms (36a, 36b), and an axle (44) which passes through the shackle (20) and the two arms. Said two arms each have an opening (48a, 48b) which is respectively provided with a first ring (50) and a second ring (60) for receiving the axle. According to the invention, the first ring has a bottom (54) which forms an abutment for a first end (44a) of the axle, and the first fixing device also comprises means for stopping the axle (64) in translation, co-operating with a second end (44b) of said axle.
92.
DEVICE FOR REINFORCEMENT OF A HOLLOW STRUCTURE, ESPECIALLY A BOX STRUCTURE FOR AN AIRCRAFT AND A HOLLOW STRUCTURE EQUIPPED WITH SUCH A DEVICE
A reinforcement device for hollow structures (1) such as aircraft housing structures, comprising a front panel (2), an upper panel (3), a lower panel (4) and a rear panel (5). The device comprises a series of reinforcing rods (6a, 6b, 6n), whose extremities are affixed to at least some of the panels (2, 3, 4, 5) comprising the structure. Some of the rods are arranged in threes so as to form a first reinforcement triangle whose peaks are affixed respectively to the front panel, essentially in a medial position on the height of the latter, as well as to the upper and lower panels. The device furthermore includes an intermediary partition placed between the front and rear panels and defining a rear housing, certain other rods are also arranged to form a second reinforcement triangle, said to be an intermediary reinforcement triangle and comprising peaks affixed respectively to the intermediary partition and to the upper and lower panels, the second triangle being located in the rear housing.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (CNRS) (France)
AIRBUS OPERATIONS SAS (France)
Inventor
Aubert, Jean-Jacques
Brot, Patrice
Crouzet, Yves
De Bonneval, Agnan
Youssef, Anis
Abstract
The invention relates to a method for the transmission of data relating to the actuation of a member used to operate a vehicle, in which code words are transmitted along the length of a given channel, each of said code words resulting from the channel coding of a block of information symbols that is representative of the aforementioned data. According to the invention, at least two such successive blocks of information symbols are coded using two different codes. The invention also relates to a data transmission system that is used to implement said method.
The process improves the maneuverability of an aircraft during the approach to landing and then flare-out phases, the aircraft being equipped with air brakes. According to the process, the air brakes are put in a first deployed position during the approach phase, and as a function of a representative parameter of a given altitude and in case of a steep angle approach, they are actuated to transition to a second more retracted position than the first position so as to achieve a flare-out allowing to essentially maintain the same angle of incidence, corresponding in case of a steep angle approach to achieve a flare- out with habitual exterior piloting references during the flare-out phase.
The invention relates to a tool (1) for crimping a cable end-contact (34) for simultaneously receiving the bare and insulated parts thereof. The inventive crimping tool comprises first and second stages for crimping said contact onto the cable conductive core and the insulating sheath thereof, respectively and means (2, 4) for substentially simultaneously actuating said first and second stages. At least one rotatable element (30) is disposed between at least one crimping element and the cam (12, 22) of at least one stage, wherein said rotatable element can take at least two positions with respect to the crimping element. A method for crimping a contact simultaneously on the insulating sheath and the conductive core thereof is also disclosed.
- The propulsion system comprises four engines (M1, M2, M3, M4) with pusher propellers, the engines being mounted in pairs on the wings (4, 5) of the aircraft (2) and the respective direction of rotation thereof when viewing said engines from the rear of the aircraft (2) being, from left to right, the clockwise direction (S1), the counterclockwise direction (S2), the clockwise direction (S1), and the counterclockwise direction (S2).
A device for extinguishing fires includes gas generation by the combustion of a pyrotechnic block connected to the means for distributing the gas generated over the fire zone. The device also includes the means to regulate the generated pressure in order to impose a profile of oxygen concentration in the fire zone. These regulation means may, for example, be a controlled valve or may be the product of the arrangement of the pyrotechnic block. The device is particularly well adapted for aircraft engine fires, since it does not use halogenated extinguishing agents.
A62C 3/08 - Fire prevention, containment or extinguishing specially adapted for particular objects or places in vehicles, e.g. in road vehicles in aircraft
A62C 37/44 - Control of fire-fighting equipment an actuating signal being generated by a sensor separate from an outlet device only the sensor being in the danger zone
98.
TRANSITION BLOCK BETWEEN THE MEANS USED TO ATTACH A WING TO AN AIRCRAFT FUSELAGE AND THE SAID WING, AND THE AIRCRAFT EQUIPPED WITH SUCH A BLOCK
A shim (20) adapted for positioning on the upper surface (13) of at least one wing (4, 5) of an aircraft. The shim (20) restores an aerodynamic profile of the wing at the position of a zone of transition (19) between structure (15) fastening the wing to the fuselage of the aircraft and the upper surface of the aircraft. The shim enables a seal to run from the fastening structure to the upper surface of the wing, in eliminating the zones at which the seal may break open. An aircraft provided with such a shim (20) is also disclosed.
The invention concerns the cabling in a connection box, using a printed circuit board which includes N input terminals on one of its edges, N output terminals on its opposite edge and between these two edges, a multitude of levels of elementary connection devices with two inputs and two outputs, the two inputs of one element, electrically connected to the two outputs of the said device, on the one hand with two direct connections and on the other hand, with two cross-linked junction boxes, the inputs of the first level devices being directly connected electrically to the input terminals, the outputs of the last level devices being directly connected electrically to the output terminals and between two successive levels, the outputs of the devices of the lower level are electrically connected the inputs of the devices of the next highest level. To configure the cabling, either the direct connections or the cross-linked connections of each device are cut.
The invention relates to a longeron (2) for an aircraft fuselage, mounted vertically between an upper surface panel (E) and a lower surface panel (I) of the fuselage and between a first wing spar (3) and a second wing spar (4 ) symmetrical to the first wing spar with relation to the fuselage, including: - peripheral zones (z1 - z3) and a central zone (c) - a recess (5) in the central zone, the peripheral zones of the longeron providing drainage of the normal forces led by the wing spars. The invention also relates to a center wing box comprising such a longeron.
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