Abstract: A method is disclosed for placing a satellite in an operational orbit. The satellite is equipped with its own satellite propulsion system as well as a detachable separate propulsion device. The satellite and separate propulsion device are launched into a transfer orbit by means of a space launcher. The separate propulsion device is controlled by a satellite. The satellite is transferred from the transfer orbit to an intermediate orbit by means of the separate propulsion device. The separate propulsion device is separated from the satellite in the intermediate orbit. The satellite then enters and operational orbit from the intermediate orbit by means of its own satellite propulsion system. The intermediate orbit is disposed between the transfer and operational orbits, and is in relatively close proximity to the operational orbit but is far enough away from the operational orbit to prevent possible interferences.

Abstract: A reinforcement insert includes superposed laps of fiber which are coupled to each other by a coupling fiber which passes through the laps in order to form fiber sections wherein at least certain sections thereof are rectilinear. The fiber sections and laps are immersed in a cured resin. The total density of the fibers that make up the insert ranges from 50% to 60% of the volume of the insert and the fiber sections take up 3%-10% of the volume of the insert. A method for creating a sandwich structure includes at least one such insert.

Abstract: A deployable wing structure for air-braking a satellite and which, once deployed, includes at least one wing structure element that forms a three-dimensional structure and includes at least two panels lying in secant planes and forming a dihedron.

Abstract: The device for thermal monitoring for a piece of equipment, whereby said piece of equipment is integrated on a craft placed in a forced vacuum environment, an outside part (7) of the equipment projecting outside of a wall (3) of the craft and being subjected to a solar radiation flux (15), comprises an absorbent screen (13) that is suitable for being placed between the outside part (7) of the equipment and the wall (3) of the craft, whereby this absorbent screen (13) has—on at least one portion of its front face, designed to be placed on the side removed from the wall of the craft—an absorptivity ?SOLAR that is the largest possible in the field of solar radiation, coupled to a low emissivity ?IR in the infrared spectrum. The absorbent screen (13) is made of a very heat-conductive material and has—on at least a portion of its rear face, suitable for being oriented toward the wall of the craft—a high emissivity ?IR in the infrared spectrum, typically greater than or equal to 0.7.

Abstract: A device for measuring the temperature of a substrate comprising a thermocouple comprising electric wires joined to each other at least one junction; a fixing element suitable for fixing said junction to said substrate in order to measure its temperature; characterized in that the fixing element comprises a thermally conductive element suitable for bearing a portion of electric wires adjacent to said junction; said thermally conductive element being capable of thermally coupling said portion of electric wires to said substrate.

Abstract: A radio antenna, particularly for a spacecraft, including a reflector and means of support of this reflector, where the reflector includes a body able to reflect radio waves, and a rigid rear structure supported by the means of support and connected to the body by decoupling angles, wherein each of said decoupling angles includes, at one at least of its ends, a layer of elastic material able to dampen at least one axial component of vibrations of the body.

Abstract: The invention concerns a radio antenna for a space satellite, including a reflector and means of support of this reflector. The reflector includes a front skin able to reflect radio waves, a rigid rear structure supported by the means of support, and a layer of elastic material interposed between said front skin and said rigid rear structure, able to dampen the vibrations of the front skin.

Abstract: Disclosed is a system for positioning a terrestrial user. The system includes navigation satellites placed in medium altitude orbits, management satellites placed in high orbits, able to manage the navigation satellites and communicate with the Earth. The management satellites can include a set of at least three management satellites each placed in a high orbit, having a plane that is inclined with respect to the plane of the terrestrial equator and cuts the equator along a diametral straight intersection line with respect to the Earth. Two external diametral straight intersection lines can form an angle of longitude of at least 90 degrees.

Abstract: A radio antenna, in particular for a spacecraft, including a reflector and a rear structure supporting said reflector, and also a rigidification membrane added on to the reflector so as to limit the displacement of a peripheral portion of the reflector in a direction parallel to a central axis of this reflector, where said rigidification membrane is separate from the rear supporting structure.

Abstract: The present disclosure relates to an altitude-adjustable apparatus comprising: a platform; a device for controllably raising the altitude of the platform; a water supply combined with said platform; a device for converting water into the components thereof through solar power; and means for liquefying and storing the resulting liquefied hydrogen in tanks, the latter being built into flying bodies that are designed to arrive at a target area and dispense the liquefied hydrogen.

Abstract: A propulsion module configured to launch a craft into space including a recoverable module and a non-recovered part, which is secured to the recoverable module at launch. The recoverable module includes a propulsive system to launch the craft, systems for command and control of the propulsive system, a subsonic flight propulsion motor, airfoils for the subsonic flight, a landing gear, and a braking parachute. The recoverable module is installed in a lower position of the craft when the craft is in the launch position. The non-recovered part includes at least one tank to feed the propulsive system. The recoverable module and the non-recovered part are configured to be separated when the propulsion module reaches a given altitude, and the recoverable module is capable of landing in a controlled fashion after a coasting flight, for example for a return to the launch site.

Abstract: A method of estimation of the direction of the line of sight comprising forming at least two images representing one and the same portion of the scene (P) at two successive instants using an instrument mounted on a moving carrier, and the determination of the change of orientation of the line of sight by matching homologous points (A1, A2) in the two images of the same portion of scene by calculating the angular variations in attitude between the two instants at which the two successive images are obtained and by correcting on the basis of the deviation, between the estimate of the point (A2) homologous in the image provided by a model, and the actual position of the homologous point, and integrating and time filtering pertaining to a sequence of several successive determinations of attitude variations, leading to an estimate of the direction of the line of sight.

Abstract: Using at least one pressure sensor in an emptying pipe or a feed pipe between emptying valves at the outlets from tanks and feed valves for feeding nozzles, to detect the pressure waves that result form opening and closing a feed valve, and then following operation of said valve, detecting the damping or disappearance of the pressure waves, indicating that the transition between the liquid propellant component and the pressurization gas is going past said pressure sensor, and thus that the tank is completely empty. This detection of complete emptying is a step in a method of managing propellant that makes it possible to allocate the remaining propellant to a residual lifetime for the satellite in orbit and/or to re-orbiting or de-orbiting the satellite.

Abstract: The gyrodyne includes an inertial wheel mounted, via a wheel support, on the moving part or rotor of a cardan assembly. The cardan assembly is provided with a stator and the rotor is rotatable with respect to the stator about a first axis of rotation, it being possible for the spinner of the inertial wheel to be set in rotation about a second axis of rotation not aligned with the first axis of rotation. The stator of the cardan is mounted on a block and fixed to this block via an arrangement of vibration attenuators or insulators. The mechanism for setting the rotor in rotation is at least partially housed in the interior volume of the block.

Abstract: The invention relates to the stiffening of inflatably deployable structures, in particular for space use. More specifically, the invention relates to a flexible membrane for inflatably deployable structures, that comprises one layer of a composite and at least one polymer film that is permeable to visible radiation covering one of the faces of this layer, said composite being formed from a fibrous material impregnated with a composition incorporating an epoxide or epoxy/acrylate resin and a photoinitiator, and which is characterized in that the photoinitiator is an iron-arene complex salt of general formula (I): where A is an arene group, while X is a non-nucleophilic anion. Applications: manufacture of devices of the following types: radars, solar panels, reflectors, solar shields, antennas, solar mirrors or sails, intended to equip orbital vehicles or vehicles for space or planetary exploration.

Abstract: An impact-resistant protective coating includes a material consisting of an outer metal adhesive layer having a sub-layer of compressible cellular material.

Abstract: A method of determining angles of sight between a receiver device having at least two reception antennas and at least one transmitter device having at least one transmission antenna for the transmission of radioelectric signals including at least two wavelengths. The determining method is based on angular phase measurements of the radioelectric signals, which are measured modulo 2? and are therefore ambiguous since they are known to within an integer number of times 2?. The ambiguity in the phase measurements is resolved in the case of an application of the determining method to a receiver device dimensioned by applying a method of design of a system for determining angles of sight. A receiver device implements the determining method, and a system for determining angles of sight includes the receiver device and at least one transmitter device.

Abstract: The thermal control device comprises at least one network of capillary heat pipes, in which each heat pipe comprises a tube enclosing an essentially annular longitudinal capillary structure, for the circulation of a two-phase heat-transfer fluid in the liquid phase, and surrounding a central channel for the circulation of said two-phase fluid in the vapor phase. The tubes of at least two heat pipes of the network intersect and are interconnected in such a way that at each intersection of heat pipes forming a node of the network, an exchange of fluid in the liquid phase can take place by capillary action between the capillary structures of said two or more heat pipes, and such that, simultaneously, an exchange of fluid in the vapor phase can take place by free circulation between the central channels of said two or more heat pipes.

Abstract: A method of using at least one polymerisable resin R1 is disclosed. According to some aspects, the resin R1 is selected from the group consisting of epoxidised polybutadiene resins, epoxidised polyisoprene resins, epoxidised polysiloxane resins, epoxidised triglyceride resins and epoxidised polyether resins. The resin R1 having at the non-polymerised state a total mass loss (TML) value lower than about 10%, a recovered mass loss (RML) value lower than about 10%, and a collected volatile condensable material (CVCM) value lower than about 1%, as determined in accordance with standard ECSS-Q-70-02A of the European Space Agency. The resin R1 also characterized by an epoxide equivalent weight of about 100 to 600 g/mole. The resin is configured for the manufacture of a composite material for use in space, and more specifically, a composite material of a Gossamer structure. Polymerisable resin compositions are also disclosed which are useful in the manufacture of composite materials for use in space.

Abstract: An electrical power supply system for a satellite with continuously controlled DC-voltage is suitable for recharging batteries without requiring any specific charge module. To this purpose, at least one solar generator section, so-called charge section, is connected directly to a battery output terminal (B). Simultaneously, at least one battery discharge module is dedicated to the control of a residual current that is not used for recharging the batteries. A first regulator is added to the system in order to control said residual current on the basis of a current produced by the charge section and a charge target for the batteries. Optionally, a second regulator may also be added, in order to directly control the current produced by the charge section when equipment that is electrically powered has a low consumption.