Abstract: The invention is related to a grid type fiber composite structure (1) comprising a grid of polygon cell modules (5) comprising at least three substantially u-shaped ribs made of fiber composite layer (8, 13) and a foam core (9) provided inside each cell module (5) for support of said u-shaped ribs. Said foam core (9) is along a base essentially in line with one of said flat cap sections (11, 12). At least one layer of strip (2, 3) is provided outside said flat cap sections (11, 12) and a skin sheet (4) is unilaterally attached to said cell modules (5) essentially in line with one of said flat cap sections (11, 12). The invention is as well related to a method of manufacturing such a grid type fiber composite structure (1).

Abstract: The present invention relates to a locking device provided with a latch (10) and a retaining pin (100), said latch (10) having a handle (11) and a locking arm (20) hinged to said handle (11) by a pivot peg (13). The retaining pin (100) comprising a first transverse branch (101) and a second transverse branch (102) separated by a space (103), a second end (22) of said locking arm (20) has a contact finger (23) provided with a first surface and with a second surface (25) opposite from said first surface (24), said first surface (24) comprising closure means for co-operating with said first transverse branch (101), said second surface (25) comprising extractor means of said latch (10) for co-operating with said second transverse branch.

Abstract: A system comprising a rotor blade (1), in particular of the tail rotor of a rotary wing aircraft, in a fiber-reinforced composite design, with a blade section (5) and with a coupling section (9) for attaching the rotor blade (1) to the hub of a drive device, and comprising a separate sleeve-shaped control tube (40) with an essentially hollow-cylindrical shaft (41), with an also essentially tubular tie section (56) to tie the control tube (40) to the rotor blade (1) by sliding it onto its coupling section (9), is improved in that the coupling section (9) of the rotor blade (1) and the tie section (56) of the control tube (40) comprise a cross-sectional shape for positive-locking interconnection of the rotor blade (1) and the control tube (40).

Abstract: A method of performing a health check of at least one turbine engine (3). During a development step (STP0), the installation losses (1) are quantified for a plurality of test values for a reduced speed of rotation (Ng?) of a gas generator (4) of the engine. During an acquisition step (STP1), the speed of rotation of said gas generator (4) is increased until said engine develops a maximum power, and then the speed of rotation of the gas generator (4) is decreased until the reduced speed of rotation (Ng?) reaches a test value. The aircraft is stabilized and at least one monitoring value is acquired. During an evaluation step (STP2) of evaluating the health check, at least one operating margin is determined by using a monitoring value and the effects of mounting the engine in an airplane.

Abstract: A method of regulating the propulsion speed of a hybrid helicopter. In application of a correction process for protecting the hybrid helicopter, piloting orders generated by a manual control member relate to an airspeed setpoint and they are corrected by control means in accordance with at least one correction mode in which the piloting orders are corrected in application of a rule that takes account of a setpoint consumed power corresponding to the airspeed setpoint derived from the piloting order.

Abstract: A method for detecting electromagnetic threats to an aircraft, wherein electromagnetic fields are measured, and in the case of an arising threat to the aircraft being detected, countermeasures are taken, characterized by the following steps: measuring the frequency and the amplitude of at least one electromagnetic field in the frequency range from 9 KHz to 40 GHz; analyzing the measured results and determining the field intensity in relation to each measured frequency; comparing the field intensity in relation to each measured frequency with stored limiting values or qualification values, wherein the aircraft is qualified for field intensities that are lower than the limiting values; and providing the result in the form of a warning signal in the case of approaching or exceeding a limiting value.

Abstract: A method of fabricating components (23) based on fiber-reinforced plastics material, in particular with the help of a wrapped mold core (10), includes the following steps: a) fabricating the mold core (10); b) wrapping the mold core (10) with a wrapper (14) that withstands resin; c) applying fiber semi-finished products (16) and introducing resin to create the plastics component (23); and d) hardening the plastics component (23) under the effect of temperature; which method is improved by the fact that, under all circumstances, and during hardening in step d), substantial pressure equilibrium is established across the fiber semi-finished product (16, 20) between ambient pressure (pu) and the pressure (pk) inside the core. A device is also described.

Abstract: A drill-elastic and flexurally rigid rod element (10) for supporting and guiding a movable flap (24) relatively to a wing (26) of an aircraft, wherein the rod element encompasses a cross-shaped profile cross section and first and second fastening sections (20, 22) and is made of a fiber composite material and the rod element can be fastened to the wing in a fixed manner via the first fastening sections (20) and to the flap (24) via the second fastening sections (22). The rod element and the fastening sections are laminated from a plurality of unidirectional, preimpregnated fibrous layers—prepreg layers—which are bonded to one another in the area of the fastening sections as well as in the cross-shaped cross section core area (28) and which are separated from one another in each case by a separating film inserted between the prepreg layers outside of the areas.

Abstract: The present invention relates to a damper (10) provided with an inner end strength member (20) and an outer end strength member (30) that are tubular and coaxial, a return member (40) being arranged between the inner end strength member (20) and the outer end strength member (30). The return member (40) includes an intermediate strength member (43) coaxial with the inner and outer end strength members (20, 30), said return member (40) having first and second return means (41, 42), said first return means (41) being arranged between the outer end strength member (30) and said intermediate strength member (43), said second return means (42) being arranged between the inner end strength member (20) and said intermediate strength member (43).

Abstract: The invention provides a vibration filter mechanism for aircraft equipment. A weighted lever arm (5) is hinged via bearings (22, 23) associated respectively with a first structure connected to a fuselage and with a second structure connected to the equipment (12). Deformable means (7) oppose pivoting movement of the lever arm. The lever arm (5) is arranged as a one-piece fork that comprises a pair of branches (19, 20) that are interconnected by a crossbar (21) and that are hinged to the bearings (22, 23) about spaced-apart parallel pivot axes (A1, A2). The fork carries a torsion shaft (24) that extends between the branches (19, 20) at their free ends (9), the torsion shaft (24) constituting the weight weighting the lever arm (5) and the deformable means (7) of the mechanism.

Abstract: The invention relates to a control system and to a method of operating such a control system, particularly to a control system for pilot command inputs for a helicopter, with a mechanical input signal (2), at least one electric position sensor (21, 21?) for said input signal (2), at least one electric power supply (25, 25?) and at least one controllable, electromechanical actuator (27, 27?) fed by the at least one electric power supply (25, 25?) and controlled by the at least one electric position sensor (21, 21?). The mechanical input signal (2) is applied mechanically to the at least one controllable electromechanical actuator (27, 27?).

Abstract: The present invention relates to an electrical power supply and control device (10) for equipment (9) of a rotor (5), the device having a three-phase AC generator (20) in a stationary frame of reference. The device (10) comprises a rotary field transformer (30) having a stationary portion (31) connected to said generator (20) and a rotary portion (32) connected to a sequencer (50) that is arranged in a rotary frame of reference, said device (10) including control means (51) for controlling the sequencer (50) and located in said stationary frame of reference, the control means being connected to the sequencer (50) via contactless connection means (200), said sequencer (50) being suitable for electrically powering said equipment (9) on order of said control means (51).

Abstract: A device for assisting the piloting of a rotorcraft, the rotorcraft having a flight control (10) for controlling a rotor (20, 21) via at least one linkage (35. An additional force is generated by at least one piloting assistance means (100) mechanically linked to the linkage (35. The additional force is a function of the position of the flight control (10) and of an instantaneous force as measured on the linkage (35).

Abstract: A fuselage structure, particularly an aircraft door (1) of composite material comprising at least one panel (2) and at least one beam (3) mounted to each other and the panel (2) with the panel (2) being formed of at least one group of composite layers (5, 6, 20). The at least one beam (3) is provided at least at one of its respective ends (10, 11) with a flange (13) suitable for adhesive engagement with the at least one panel (2). The at least one group of composite layers (5, 6, 20) of the panel (2) is in form locking engagement with this flange (13) of the beam (3).

Abstract: The present invention relates to a rotorcraft blade (1) provided with at least one spar (3) and an outer covering (4), said blade (1) including a hollow sleeve (10) with a wall (11) that is provided at least with said outer covering (4) secured to said spar (3). The blade includes a rotary damper (20) provided with an inner tubular strength member (21) and an outer tubular strength member (22) that are coaxial and connected together by a ring (23) of viscoelastic material, said inner tubular strength member (21) being secured to said covering (4), and said outer tubular strength member (22) being provided with a fork (25) possessing first and second tines (26, 27) facing each other so as to be suitable for receiving a first ball joint (41) close to the pitch-variation axis (AX1) of said blade (1).

Abstract: A method of determining the locations of loads in a vehicle in such a manner as to ensure mission safety in terms of loads being retained in the aircraft. On the basis of vehicle parameters, the anchor points available in this type of vehicle for retaining the loads are identified. As a function of the available loading volume(s), a possible loading configuration is calculated using at least some of the available anchor points, at least some of the retaining means, and the characteristics of the loads. The forces that might be exerted on the selected anchor points are verified as being less than maximum limits. The calculated loading configuration is confirmed, otherwise a new loading configuration is calculated.

Abstract: A method for producing a continuous, three-dimensional, closed semi-finished product made of fiber composite from at least one planar, flat fibrous preform. The method is characterized in that the flat fibrous preform is initially placed flat and aligned to produce a load-bearing fiber orientation and subsequently formed into a three-dimensional structure. In a final step, the three-dimensional structure is then closed to form a closed semi-finished product.

Abstract: A rotary aircraft which carries a hanging sling load on at least one support cable, benefits from a stabilizer which generates control signals R to be applied or modulated onto the flight control of the aircraft. At least one sling load pendulum motion detecting unit provides the stabilizer with data on the pendulum motion of the sling load. Additionally, the stabilizer is set up with a working range determination unit S3 for determining the space where free movement is possible in which the rotary aircraft with sling load can move without collision. The stabilizer generates control signals R in dependency of the determined pendulum motion of the sling load and the determined working range.

Abstract: A device (10) for damping the vibration of a structure is provided. The device includes a main mass (12) and an auxiliary mass (20) mounted on a support (11) that is suitable for having one end held fixedly in the structure (1). The auxiliary mass (20) includes a threaded member (21) arranged around a drive screw (40) inside a casing (30) fastened to the main mass (12). The device includes manual drive (50) engaging the drive screw (40) to move the auxiliary mass (20) along the drive screw (40).

Abstract: An anti-crash seat (1) for a vehicle, the seat comprising a bucket (2) and a frame (6), said bucket (2) being provided with a seat proper (3) and with a back (4) for receiving a person. Said seat (1) includes crash energy absorber means (10) and antivibration means (20). The seat also includes inhibitor means (30) for inhibiting said antivibration means (20) so as to neutralize the antivibration means (20) when the seat (1) is subjected to acceleration towards the floor that is greater than a predetermined threshold.