Abstract: An airship is equipped with a compact solar generator using concentration to supply the airship in flight with electrical energy from solar radiation. The compact solar generator comprises a first set of row(s) of bifacial photovoltaic solar cells, arranged parallel to a longitudinal central axis of the airship, and a solar radiation concentrator for making solar rays converge towards rear faces of the bifacial solar cells of the first set. The solar radiation concentrator is a second set of one or more local solar radiation concentrator(s), wherein each local concentrator is paired with a corresponding row of solar cells and comprises a reflector of convex form suitable for making solar radiation converge towards the rear faces of the solar cells of the paired row.

Abstract: A method of guidance for placing a satellite on station comprises the following steps carried out during a predefined current cycle: A) determining on the ground a law of orientation of the thrust vector, and a history of state variables and of adjoint state variables of the satellite for the transfer from a starting orbit to a target orbit using optimal control theory, B) determining on the ground a law of evolution of the rotation of the satellite about the thrust vector, on the basis of the orientation law and of the history, C) representing according to a predetermined format the evolution of the state variables and adjoint state variables so as to obtain first parameters, D) representing according to a predetermined format a law of evolution of the rotation so as to obtain second parameters, E) concatenating the first and second parameters so as to obtain a guidance plan for the satellite, F) downloading onboard the guidance plan, G) periodically repeating according to a predefined period which is smalle

Abstract: A balloon comprises an envelope containing a lifting gas and a concentrated solar radiation solar generator. The solar generator includes a reflector, one or two arrays of photovoltaic solar cells forming a first active face directed towards the reflector and a second active face directed towards the exterior of the envelope of the balloon. The reflector, the first active face and the second active face of the array of photovoltaic cells are configured so as to ensure the first active face and the second active face of the array both generate electrical power provided that the rollwise solar misalignment of the reflector is smaller than or equal to 10 degrees in absolute value.

Abstract: A method for stationing a satellite, comprises: determining a predefined trajectory for stationing of the satellite based on a model of movement of the satellite; determining parameters of the predefined control law of approximation of the trajectory and by minimizing impact on the control law of deviation from a trajectory followed by the satellite using parameters of the control law; determining a state vector of the satellite; determining a deviation between the state vector and the predefined trajectory; determining Lagrange multipliers based on a current state vector of the satellite, on a deviation between current state vector and predefined trajectory and on parameters of the predefined control law; determining parameters of the current control law of the engines based on the Lagrange multipliers and by derivation of a parameter representative of an effect of engines on the real trajectory; and controlling engines based on parameters of the current control law.

Abstract: A method of guidance for placing a satellite on station comprises the following steps carried out during a predefined current cycle: A) determining on the ground a law of orientation of the thrust vector, and a history of state variables and of adjoint state variables of the satellite for the transfer from a starting orbit to a target orbit using optimal control theory, B) determining on the ground a law of evolution of the rotation of the satellite about the thrust vector, on the basis of the orientation law and of the history, C) representing according to a predetermined format the evolution of the state variables and adjoint state variables so as to obtain first parameters, D) representing according to a predetermined format a law of evolution of the rotation so as to obtain second parameters, E) concatenating the first and second parameters so as to obtain a guidance plan for the satellite, F) downloading onboard the guidance plan, G) periodically repeating according to a predefined period which is smalle

Abstract: A method for stationing a satellite, comprises: determining a predefined trajectory for stationing of the satellite based on a model of movement of the satellite; determining parameters of the predefined control law of approximation of the trajectory and by minimizing impact on the control law of deviation from a trajectory followed by the satellite using parameters of the control law; determining a state vector of the satellite; determining a deviation between the state vector and the predefined trajectory; determining Lagrange multipliers based on a current state vector of the satellite, on a deviation between current state vector and predefined trajectory and on parameters of the predefined control law; determining parameters of the current control law of the engines based on the Lagrange multipliers and by derivation of a parameter representative of an effect of engines on the real trajectory; and controlling engines based on parameters of the current control law.

Abstract: A solar panel (3) having a front face (31) for receiving solar radiation and a rear face (32), for use in a spacecraft (1). The solar panel is covered on the rear face (32) with a coating (4) having negative electric charge emissive property such that in use the coating expels electrons out of the rear face (32) of the solar panel (3), thus reducing the potential difference between the front face (31) and the rear face (32). A method of reducing potential difference between the front face (31) and the rear face (32) of the solar panel (3) by covering the rear face by said coating (4).

Abstract: The invention relates to a thermal control device intended to dissipate the heat generated by a payload on a spacecraft, comprising a number of surfaces and including means for circulating a refrigerant. An evaporation zone (Z1) comprises means for circulating the refrigerant, a compression zone (Z2), a condensation zone (Z3) comprising at least one radiating panel, linked to a part of the means for circulating the refrigerant, including several branches and comprising means to allow or inhibit the circulation of the refrigerant within these branches so as to vary the area of the heat exchange surface in the condensation zone, a pressure reduction zone (Z4) comprising means for circulating the refrigerant. Such a device is particularly well adapted to thermal problems encountered in telecommunications satellites.

Abstract: A structural panel for a satellite comprises an outer skin intended to be outside the satellite, a core comprising at least one integrated heat pipe mounted in fixed contact with said outer skin, and an inner skin intended to be inside the satellite, said structural panel being equipped with generic heat exchangers which are adapted to be associated with a heat control circuit using the circulation of a liquid coolant, said circuit being situated outside the panel.

Abstract: A system is dedicated to the control of the deployment of at least two spacecraft (ES1, ES2) which are provided with maneuvering means (MD1, MD2) and are intended to move according to a chosen formation. This system includes a control device (MM11, MM12, MT, MC1) comprising i) first measurement means (MM11, MM12, MT) responsible for determining substantially simultaneously and with high precision the orbital positions of the spacecraft (ES1, ES2), and ii) first calculation means (MC1) responsible for determining for each of the spacecraft, as a function of their orbital positions, maneuvers intended to position each of them at a chosen instant substantially in a chosen position with respect to a reference trajectory, having regard to the time law of a reference craft (ES1) on this reference trajectory (TR), so as to place the formation in a chosen configuration.

Abstract: The invention produces contactless radial position sensors having an equivalent angular aperture greater than 90°, and substantially eliminates the spurious microvibrations due to the imperfections of the current sensors. For this, the invention uses at least four independent sensors and an electronic device capable of virtually placing in series said sensors in groups of two or three in order to spread the equivalent angular aperture of said groups of sensors.

Abstract: The invention relates to a thermal management device intended to collect the heat dissipated by a group of dissipative equipment of a spacecraft in an evaporation zone before discharging this heat to cold space via a condensation zone, situated in the region of the radiators. The evaporation zone and the condensation zone each consist of at least one heat-exchange part comprising a network of a plurality of compact heat-exchange elements distributed over their respective heat-exchange surface and connected in series and/or in parallel by the tubes of the thermal fluid circulation means. A device such as this is particularly intended for telecommunications satellites.

Abstract: A system is dedicated to the control of the deployment of at least two spacecraft (ES1, ES2) which are provided with maneuvering means (MD1, MD2) and are intended to move according to a chosen formation. This system includes a control device (MM11, MM12, MT, MC1) comprising i) first measurement means (MM11, MM12, MT) responsible for determining substantially simultaneously and with high precision the orbital positions of the spacecraft (ES1, ES2), and ii) first calculation means (MC1) responsible for determining for each of the spacecraft, as a function of their orbital positions, maneuvers intended to position each of them at a chosen instant substantially in a chosen position with respect to a reference trajectory, having regard to the time law of a reference craft (ES1) on this reference trajectory (TR), so as to place the formation in a chosen configuration.

Abstract: The invention produces contactless radial position sensors having an equivalent angular aperture greater than 90°, and substantially eliminates the spurious microvibrations due to the imperfections of the current sensors. For this, the invention uses at least four independent sensors and an electronic device capable of virtually placing in series said sensors in groups of two or three in order to spread the equivalent angular aperture of said groups of sensors.

Abstract: The invention relates to a thermal control device intended to dissipate the heat generated by a payload on a spacecraft, comprising a number of surfaces and including means for circulating a refrigerant. An evaporation zone (Z1) comprises means for circulating the refrigerant, a compression zone (Z2), a condensation zone (Z3) comprising at lEast one radiating panel, linked to a part of the means for circulating the refrigerant, including several branches and comprising means to allow or inhibit the circulation of the refrigerant within these branches so as to vary the area of the heat exchange surface in the condensation zone, a pressure reduction zone (Z4) comprising means for circulating the refrigerant. Such a device is particularly well adapted to thermal problems encountered in telecommunications satellites.

Abstract: A device is disclosed for reducing the risk of primary and secondary electrostatic discharges occurring in particular in the solar generators of spacecraft. A solar energy concentrator device for spacecraft includes a reflector for reflecting solar radiation onto a photovoltaic cell for converting solar energy into electrical energy and a heat transfer arrangement for transporting to a cold area heat energy stored by the cell following reception of solar radiation. One particular application of the device is to solar panels for satellites.

Abstract: An action device comprises first electromagnetic means installed on a first body and defining a first magnetic moment and a magnetic field and second electromagnetic means installed on a second body, distant from the first body, and defining a second magnetic moment able to interact with the magnetic field. It further comprises i) means for varying the first magnetic moment in accordance with a chosen first law of variation to cause the magnetic field to vary in time, ii) means for varying the second magnetic moment in accordance with a second law of variation so that a required force and a required torque are induced on the second body, and iii) calculation means adapted to determine the second law of variation as a function at least of the required force and the required torque and the first law of variation.

Abstract: A solar panel (3) having a front face (31) for receiving solar radiation and a rear face (32), for use in a spacecraft (1). The solar panel is covered on the rear face (32) with a coating (4) having negative electric charge emissive property such that in use the coating expels electrons out of the rear face (32) of the solar panel (3), thus reducing the potential difference between the front face (31) and the rear face (32). A method of reducing potential difference between the front face (31) and the rear face (32) of the solar panel (3) by covering the rear face by said coating (4).

Abstract: A device is disclosed for reducing the risk of primary and secondary electrostatic discharges occurring in particular in the solar generators of spacecraft. A solar energy concentrator device for spacecraft includes a reflector for reflecting solar radiation onto a photovoltaic cell for converting solar energy into electrical energy and a heat transfer arrangement for transporting to a cold area heat energy stored by the cell following reception of solar radiation. One particular application of the device is to solar panels for satellites.

Abstract: The invention provides a momentum wheel energy storage device for a spacecraft or satellite. The device has six wheels whose axes of rotation are parallel to edges of a tetrahedron. It thus provides redundancy which makes it possible simultaneously to store energy and to generate an arbitrary net angular momentum, even in the event of one of the wheels failing. The failure of one wheel can be accommodated without degrading operation, with energy storage being optimal regardless of the angular momentum to be generated. The failure of two wheels can be accommodated but operation may then be in degraded mode. The invention makes it possible to limit the number of momentum wheels used in a device that can accommodate failures.