Patent Publication Number: US-10773832-B2

Title: Projectile intended for damping a spacecraft and corresponding space delivery vehicle

Description:
TECHNICAL FIELD 
     The present invention concerns the field of attitude control in spacecraft, such as satellites. The disclosure relates more particularly to a spacecraft comprising active attitude control means and passive attitude control means. One particularly advantageous but non-limiting application of the invention is the case of satellites in a low orbit. 
     PRIOR ART 
     The orbit and attitude control system of a spacecraft is essential to the correct performance of a mission, regardless of the type of spacecraft concerned (satellite, space platform, launch vehicle, deep-space probe, etc.) 
     The term “attitude control” used herein refers to the control of the spacecraft&#39;s orientation, i.e. generally speaking, of the movement of said spacecraft about the centre of mass thereof (also known as the centre of inertia). 
     To date, when the spacecraft is operating in an operational orbit, different means exist for actively controlling the attitude thereof, which means use electronics, information technology and sensors and actuators which consume energy and have a limited life span. For example, in the case of a three-axis stabilised satellite, active attitude control means are known, such as:
         inertial actuators that do not change the total angular moment of the satellite, such as reaction wheels or gyroscopic actuators,   actuators that do change the total angular moment of the satellite, such as thrusters or magnetic torque rods.       

     When the active attitude control means stop working, either as a result of a malfunction or because the power sources supplying them are spent or have malfunctioned, said spacecraft is considered to have become space debris. Such debris is known to be especially animated by a rotational movement about the centre of mass thereof, sometimes with a high rotational speed, for example of several degrees per second. Multiple causes of such a rotational speed can exist: accumulated internal moment, thruster malfunction, external torque caused by the solar radiation pressure, etc. In a manner known by a person skilled in the art, those debris have a variable, arbitrary attitude and are typically animated by a Poinsot motion. 
     The presence in space of such space debris is problematic since it creates space pollution insofar as the latter follow trajectories that can cross those occupied by functional spacecrafts, which creates collision risks. Moreover, collisions between debris increase the total number of debris particles, which further accentuates the risk of collision for functional spacecrafts. 
     Said space debris can also cross the orbits of other functional satellites, which creates collision risks. 
     Moreover, in order to capture and remove such space debris from orbit, other spacecrafts are known and are suitable for performing manoeuvres such as docking onto debris, in order to form a composite, like for example deorbiting satellites such as those disclosed in patent applications EP 2746163 and EP 2671804. Nonetheless, it is understood that the rotational speed of the debris remains a factor that limits the success of such a capture/deorbiting mission. More specifically, the higher the rotational speed, the more said manoeuvres, in particular capture, are difficult to achieve. Moreover, even in the case of a successful capture operation, the subsequent operations for controlling the composite and removing it from orbit are incompatible with a high rotational speed, in particular when the debris is connected to the deorbiting satellite by flexible ties, for example a harpoon. Slowing of the rotational speed of the debris is thus sought for the successful deorbiting thereof. 
     SUMMARY 
     The purpose of the present disclosure is to overcome all or part of the drawbacks of the prior art, in particular those described hereinabove, by proposing a solution that produces a spacecraft comprising active attitude control means, in addition to passive attitude control means suitable for slowing the rotation of said spacecraft when the active control means thereof have developed a permanent malfunction. 
     For this purpose, and according to a first feature, the disclosure relates to a spacecraft comprising a main body and an attitude control system, said attitude control system comprising active attitude control means suitable for the 3-axis stabilisation of the attitude of said spacecraft. Moreover, said attitude control system further comprises passive attitude control means suitable for generating, in cooperation with the Earth&#39;s magnetic field, a damping torque and comprising at least one passive damper, said at least one passive damper comprising an outer enclosure and an inner body that are configured such that:
         said inner body is positioned inside said outer enclosure and is capable of moving in rotation inside said outer enclosure about at least one axis of rotation,   said outer enclosure comprises an inner surface and said inner body comprises an outer surface, said surfaces being separated by means of a viscous fluid,   said inner body is permanently magnetised,   said outer enclosure is fixed to the main body of said spacecraft for rotation therewith.       

     According to specific embodiments, the spacecraft can further comprise one or more of the following features, which must be considered singly or according to any combinations technically possible. 
     In one specific embodiment, the active attitude control means are at least one reaction wheel or at least one gyroscopic actuator. 
     In one specific embodiment, the active attitude control means generate an active control torque of maximum value C a , and the passive control means generate a damping torque of maximum value C p  such that the ratio C a /C p  is substantially greater than 10, preferably substantially greater than 100. 
     In one specific embodiment, the passive control means of the spacecraft, intended to be subjected to external perturbation torque, generate a damping torque of maximum value C p  that is less than said external perturbation torque by a factor of substantially equal to 10, preferably substantially equal to 100. 
     In one specific embodiment, said main body comprises an inner surface, said at least one passive damper being positioned on said inner surface. 
     In one specific embodiment, said main body comprises an outer surface, said at least one passive damper being positioned on said outer surface. 
     In one specific embodiment, said at least one passive damper is devoid of any means for maintaining a gap between said outer enclosure and said inner body. 
     In one specific embodiment, said inner body is capable of moving in rotation inside said outer enclosure about arbitrary axes of rotation. 
     In one specific embodiment, the inner surface of the outer enclosure and the outer surface of the inner body are respectively spherical in shape. 
     In one specific embodiment, said spacecraft comprises a plurality of passive dampers. 
     In one specific embodiment, said inner body is capable of moving in rotation inside said outer enclosure about a single axis of rotation. 
     In one specific embodiment, the inner surface of the outer enclosure and the outer surface of the inner body each have a single axis of revolution that is substantially aligned with said single axis of rotation, and are furthermore substantially proportional. Surprisingly, the use of a single passive damper in such an arrangement produces a deceleration. 
     In one specific embodiment, the inner surface of the outer enclosure and the outer surface of the inner body are respectively cylindrical in shape. 
     In one specific embodiment, said spacecraft comprises two passive dampers arranged such that the axes of rotation of the respective inner bodies thereof are not parallel in pairs. 
     In one specific embodiment, said spacecraft comprises three passive dampers arranged such that the axes of rotation of the respective inner bodies thereof are not parallel in pairs. 
     In one specific embodiment, the inner body comprises an inner enclosure comprising a permanent magnet fixed to said inner enclosure for rotation therewith. 
     In one specific embodiment, said inner enclosure comprises polyurethane foam inside a volume that is not occupied by said permanent magnet. 
     In one specific embodiment, the inner body is solid. 
     In one specific embodiment, at least one of either the inner surface of the outer enclosure or the outer surface of the inner body is made of a non-metallic material. 
     In one specific embodiment, the inner surface of the outer enclosure is made of a synthetic resin of the polyimide type. 
     In one specific embodiment, the densities respectively of the inner body and of the viscous fluid are substantially equal for at least one temperature of said viscous fluid lying in the interval [10° C., 30° C.]. 
     In one specific embodiment, the pressure of the viscous fluid is substantially equal to 5 bar for at least one temperature lying in the interval [10° C., 30° C.]. 
     In one specific embodiment, the viscous fluid is silicon oil of the dimethicone type. 
     According to a second feature, the present disclosure relates to a spacecraft comprising a main body and an attitude control system, said attitude control system comprising active attitude control means suitable for the 3-axis stabilisation of the attitude of said spacecraft by generating an active control torque of maximum value C a . Moreover, said attitude control system further comprises passive attitude control means suitable for generating, in cooperation with the Earth&#39;s magnetic field, a damping torque of maximum value C p  such that the ratio C a /C p  is substantially greater than 10, preferably substantially greater than 100. 
     In specific embodiments, the passive control means of the spacecraft, intended to be subjected to external perturbation torque, generate a damping torque of maximum value C p  that is less than said external perturbation torque by a factor of substantially equal to 10, preferably substantially equal to 100. 
     According to a third feature, the present disclosure relates to a projectile intended for damping a spacecraft comprising a main body and active attitude control means suitable for the 3-axis stabilisation of the attitude of said spacecraft, said projectile comprising a harpoon and being intended to equip a space delivery vehicle, that is different from the spacecraft, in order to be projected from said space delivery vehicle towards said spacecraft. Moreover, said projectile comprises a passive damper mounted such that it is fixed on said harpoon and suitable for generating, in cooperation with the Earth&#39;s magnetic field, a damping torque, said passive damper comprising an outer enclosure and an inner body that are configured such that:
         said inner body is positioned inside said outer enclosure and is capable of moving in rotation inside said outer enclosure about at least one axis of rotation,   said outer enclosure comprises an inner surface and said inner body comprises an outer surface, said surfaces being separated by means of a viscous fluid,   said inner body is permanently magnetised,   said outer enclosure is fixed to the main body of said spacecraft for rotation therewith once the harpoon is secured to said main body.       

     In specific embodiments, the inner body of the passive damper mounted such that it is fixed on the harpoon is capable of moving in rotation inside said outer enclosure about a single axis of rotation, the inner surface of the outer enclosure and the outer surface of the inner body each having a single axis of revolution that is substantially aligned with said single axis of rotation, and which are furthermore substantially proportional. 
     According to a fourth feature, the present disclosure relates to a space delivery vehicle comprising at least one projectile according to the disclosure. 
    
    
     
       PRESENTATION OF THE FIGURES 
       The characteristics and advantages of the invention will be better understood after reading the following description, which discloses particular embodiments, which are in no way limiting. 
       The description is given with reference to the accompanying figures, which show: 
         FIG. 1 : a diagrammatic view of an example embodiment of a spacecraft comprising active attitude control means and passive attitude control means; 
         FIG. 2 : a diagrammatic view of one specific embodiment of the passive control means in  FIG. 1 , wherein said passive control means comprise at least one passive damper comprising an inner body capable of moving in rotation about arbitrary axes of rotation; 
         FIGS. 3 and 4 : diagrammatic views of alternative embodiments of said at least one passive damper in  FIG. 2 , wherein said inner body is capable of moving in rotation about a single axis of rotation; 
         FIG. 5 : a diagrammatic view of a preferred alternative embodiment of the spacecraft in  FIG. 1 , wherein said spacecraft comprises three passive dampers as shown in  FIG. 3 ; and 
         FIG. 6 : a diagrammatic view of an example of a projectile attached to a space delivery vehicle. 
     
    
    
     In these figures, identical reference numerals in one or another figure denote identical or similar elements. For clarity purposes, the elements shown are not to scale, unless specified otherwise. 
     DETAILED DESCRIPTION OF EMBODIMENTS 
     The present invention relates to the field of passive attitude control of a spacecraft rotating about itself. 
       FIG. 1  diagrammatically illustrates one example embodiment of a spacecraft  100  comprising an attitude control system. 
     The term “attitude control” used herein refers to the control of the orientation of the spacecraft  100 , i.e. generally speaking, of the movement of said spacecraft  100  about the centre of mass thereof (also known as the centre of inertia). This orientation can more particularly correspond to the direction of a predetermined axis defined in a frame of reference associated with the geometrical configuration of said spacecraft  100 . Moreover and more specifically, the term “active” (and respectively “passive”) is used to describe an attitude control, the implementation whereof consumes at least one (and respectively does not consume any) electrical or chemical power source on-board the spacecraft  100 . 
     Moreover, said spacecraft  100  comprises orbit control means that are not shown in  FIG. 1 . The term “orbit control” is used herein to refer to the control of the movement of the centre of mass of said spacecraft  100  in an inertial frame of reference. For example, said orbit control means are chemical and/or electrical thrusters. 
     The description provided below more specifically, however in a non-limiting manner, concerns the case in which the spacecraft is a satellite  100 , intended to rotate about itself in space in an orbit such that the magnetic field affecting this orbit allows for the operation of the invention. Preferably, the satellite  100  was initially placed in a low Earth orbit, also known by a person skilled in the art as “LEO”, by means of a launch vehicle, and is currently in a spent configuration in its final orbit. The term “low Earth orbit” is used herein to refer to an orbit that is typically less than 3,000 km, or even less than 1,500 km, whereby the Earth&#39;s magnetic field affecting this type of orbit is about 30 μT at 1,000 km. The terms “spent configuration” are used herein to refer to a satellite that is unable to continue the mission assigned thereto, for example as a result of a malfunction affecting the attitude control system, so much so that it is considered to have become space debris. Moreover, the inclination of the orbit of the satellite  100  is sufficient to allow the Earth&#39;s magnetic field to have a variable orientation in a local orbital frame of reference. For example, the satellite  100  advantageously follows a polar or sun-synchronous orbit. 
     However, according to other examples not described herein, other types of spacecraft (space platform, launch vehicle, deep-space probe, etc.) can also be considered. The disclosure can thus apply in a more general manner to any spacecraft in a low orbit, the rotation whereof about itself is sought to be passively controlled when in a spent configuration. 
     The satellite  100  comprises a main body  110 . In practice, said main body  110  encompasses a certain number of common equipment, not shown in  FIG. 1 , such as motors, sensors and miscellaneous electronics, etc. Moreover, the main body  110  of the satellite  100  can also be combined with other elements so as to form the entire satellite  100 , said other elements not being shown in  FIG. 1  since, as such, they do not fall within the scope of the disclosure. For example, said other elements are solar panels arranged on either side of the main body  110 , secured to the main body  110 . 
     In the non-limiting example shown in  FIG. 1 , the main body  110  of the satellite  100  has a substantially cubic shape. However, other shapes of the main body  110 , such as cylindrical, spherical, parallelepipedal and rectangular shapes, etc. are also possible. 
     Said main body  110  comprises an outer surface  111  that is closed and intended to be exposed to a vacuum, as well as an inner surface  112  that is closed and opposite said outer surface  111 . 
     The satellite  100  comprises an attitude control system comprising, on the one hand, active attitude control means  150 . Said active attitude control means are suitable for the three-axis stabilisation of the attitude of said satellite  100  by generating an active control torque of maximum value C a . 
     Preferably, said active attitude control means  150  comprise actuators such as momentum wheels (reaction wheels, gyroscopic actuators) or magnetic torque rods. For example, and as shown in  FIG. 1  in a non-limiting manner, said active attitude control means comprise three reaction wheels  150  respectively generating three auxiliary active control torques, the respective directions of these auxiliary torques being linearly independent such that the sum thereof forms said active control torque. 
     On the other hand, the attitude control system comprises passive attitude control means  200 . Said passive attitude control means  200  are suitable for generating, in cooperation with the Earth&#39;s magnetic field, a damping torque so as to slow the rotation of the satellite  100  about itself. 
     In one specific embodiment, said passive control means  200  generate a damping torque of maximum value C p  such that the ratio C p /C a  is substantially greater than 10, preferably substantially greater than 100. Such a torque ratio C p /C a  is advantageous in that, when the satellite  100  is on a mission, it makes any effect caused by the operation of the passive attitude control means  200  on the operation of the active attitude control means  150  negligible. More specifically, since it does not consume power, it is understood that said passive control means  200  generate a damping torque both during a mission and in the spent configuration. Thus, the term “negligible” is understood herein to refer to an effect that is low enough to ensure that the use of the passive control means  200  does not require, in order to correctly perform the mission of the satellite  100 , the use of active control means  150  that are oversized (in particular in terms of weight, bulk and cost) relative to a standard configuration that does not comprise passive control means  200 . 
     Moreover, the satellite  100  is subjected, in its orbit, in a manner known to a person skilled in the art, to external perturbation torques, for example atmospheric drag, solar radiation pressure, and a gravity gradient, etc. Thus, in one specific embodiment of the invention, the passive control means  200  are configured so as to generate a passive control torque of maximum value C p  that is less than said external perturbation torques by a factor of substantially equal to 10, preferably substantially equal to 100. Insofar as the active control means are sized such that they correct the attitude of the satellite  100  facing those external perturbation torques, it is understood that the effect of the passive control means on the active control means remains negligible in such a configuration. 
     However, it is understood that the arrangements described hereinabove and relative to the ratio between the maximum torque of the passive control means  200  and, either the maximum torque of the active control means  150  or the external perturbation torque, constitute additional characteristics. Thus, an intended purpose, i.e. producing a satellite  100  comprising active attitude control means  150 , and passive attitude control means  200  suitable for slowing the rotation of said spacecraft  100 , can be achieved without the need to employ these additional characteristics. 
     As illustrated in  FIG. 6 , it should be noted that the passive control means  200  is a passive damper that can be delivered to the satellite  100  by another space vehicle  300 , for example when the attitude control system of said satellite  100  has malfunctioned. In such a case, a so-called “space delivery vehicle” as described in patent application EP2746163, could approach the satellite  100  and deploy a harpoon  202 , on which such a passive damper  200  is fixedly mounted, so as to form a projectile  210  intended for damping the satellite  100 . The harpoon  202  would, for example, be adapted from the harpoon described in the patent application EP2671804 in order to carry the passive damper. The term “fixedly mounted” is used herein to refer to the fact that the passive damper  200  engages with the harpoon  202  in a fixed manner, such that it is rigidly secured thereto, such engagement being possible in any area of the harpoon. For example, the passive damper  200  is fixedly mounted on one of the ends of said harpoon, or between the ends of said harpoon  202  on the outer surface thereof. Once projected, such a harpoon  202  equipped with the passive damper  200  does not retain any physical ties with the space delivery vehicle  300 . 
     Thus, when the satellite  100  is in the spent configuration thereof, and thus capable of being animated by a high rotational speed, the space delivery vehicle firstly enters a phase in which it approaches the satellite  100 . When the distance separating it from the satellite  100  is equal to a predetermined distance, for example calculated in advance by means of digital simulations, the space delivery vehicle adopts an appropriate attitude ensuring that the projectile is pointing towards the satellite  100 . The projectile is then deployed towards the satellite  100  in order to reach same and thus secure the harpoon, via one of the ends thereof, to the body  110  of the satellite  100 . The rotation of the satellite  100 , which is not mechanically connected to the delivery vehicle, is then dampened thanks to the passive damper  200  before said satellite is captured in order to be removed from the trajectory thereof. Capture of the satellite  100  is, for example, performed by said space delivery vehicle, however other known means of capture are also possible. 
     A person skilled in the art will clearly see that one possible alternative design of the space delivery vehicle consists in considering same to comprise, not just a single projectile as described hereinabove, but a plurality of projectiles, for example three, said projectiles being intended to reach different points of the body  110  of the satellite  100 . For this purpose, the projectiles preferably reach the satellite  100  such that they are spaced apart by a predetermined minimum distance, in order to minimise any possible magnetic perturbation between the respective inner bodies  220  of the passive dampers  200 . A person skilled in the art knows how to determine such a distance, since the spatial decay of the electromagnetic field generated by an inner body  220  follows a known law. 
       FIG. 2  diagrammatically illustrates one specific embodiment of the passive control means  200  in  FIG. 1 , wherein said passive control means  200  comprise at least one passive damper  200 , a sectional view whereof is shown in  FIG. 2 . 
     In the non-limiting example shown in  FIG. 2 , said passive damper  200  comprises an outer enclosure  210  and an inner body  220 . 
     As shown in  FIG. 2 , the inner body  220  is positioned inside said outer enclosure  210 . Moreover, the outer enclosure  210  comprises an inner surface  211  that is closed, and the inner body  220  comprises an outer surface  221  that is closed, and that is opposite and separate from the inner surface  211  of the outer enclosure  210 . The inner surface  211  and outer surface  221  respectively of the outer enclosure  210  and of the inner body  220  are separated by means of a viscous fluid  230 . 
     In one preferred embodiment, shown in a non-limiting manner in  FIG. 2 , the passive damper  200  is devoid of any means for maintaining a gap between the outer enclosure  210  and the inner body  220 . Such a configuration is advantageous since it simplifies the production of the passive damper  200 , and allows the inner body  220  to move freely in the volume delimited by the inner surface  211  of the outer enclosure  210 . The term “freely” is understood herein to mean that the outer surface  221  of the inner body  220  is capable of being in contact with the inner surface  211  of the outer enclosure  210 . 
     However, according to other examples not described herein, the passive damper  200  can also comprise means for maintaining a gap between the outer enclosure  210  and the inner body  220 . For example, said gap-maintaining means comprise ball bearings, evenly distributed around the inner body  220  in the space occupied by the viscous fluid. 
     Moreover, the outer enclosure  210  of the passive damper  200  is fixed to the main body  110  of the satellite  100  for rotation therewith. For example, the outer enclosure  210  comprises an outer surface  212  that is closed and opposite the inner surface  211  of the outer enclosure  210 , said outer surface  212  being maintained fixed to the main body  110  by bonding, welding or any other means. 
     In one specific embodiment, and as shown in  FIG. 1 , said at least one passive damper  200  is positioned on the inner surface  112  of the main body  110 . 
     In another specific embodiment, said at least one passive damper  200  is positioned on the outer surface  111  of the main body  110 . 
     The passive damper  200  is also designed such that the inner body  220  is capable of moving in rotation inside the outer enclosure  210  about at least one axis of rotation. 
     In one specific embodiment, said inner body  220  is capable of moving in rotation inside the outer enclosure  210  about arbitrary axes of rotation. In this manner, the freedom of movement of the inner body  220  inside the outer enclosure  210  is virtually total since this rotational mobility of the inner body  220  is cumulated with the fact that no mechanical device opposes the movement of said inner body  220  until contact is made with the outer enclosure  210 . 
     For example, and as shown in  FIG. 2  in a non-limiting manner, the inner surface  211  of the outer enclosure  210  and the outer surface  221  of the inner body  220  preferably each have a spherical shape. Such a configuration is advantageously suited for rotational mobility about arbitrary axes. In the case whereby said inner surface  211  and outer surface  221  come into contact with one another, said contact is limited, at all times, to a restricted area of said surfaces so as to minimise the risks of attachment resulting from the Van der Waals forces between these two surfaces. As regards these attachment risks, the choice of materials is also important, as explained hereafter. However, the inner surface  211  of the outer enclosure  210  and the outer surface  221  of the inner body  220  can also take on other shapes, for example respectively cylindrical shapes, as described hereafter in an alternative embodiment. Moreover, the satellite  100  can also comprise a plurality of passive dampers  200 , that are respectively capable of moving in rotation inside the outer enclosure  210  about arbitrary axes of rotation. 
     The passive damper  200  is furthermore configured such that said inner body  220  is permanently magnetised. Thus, when the satellite  100  is on a mission travelling about the Earth, the inner body  220  is subjected to a magnetic torque as a result of the interaction thereof with the Earth&#39;s magnetic field so as to align the magnetic moment of the inner body  220  with the Earth&#39;s magnetic field. This interaction is independent from the operating state of the satellite  100 , and continues to exist when the latter is in the spent configuration thereof. 
     More particularly, the magnetic torque to which the inner body  220  is subjected is a function of the respective intensities of the Earth&#39;s magnetic field and of the magnetic moment of the inner body  220 . Thus, in the present embodiment, the satellite  100  is orbiting such that the Earth&#39;s magnetic field has a sufficient intensity and the dimensions of the inner body  220  are equal to about several centimetres. For example, the outer surface  221  of the inner body  220  is a sphere, the radius whereof is substantially equal to 5 cm. 
     In a preferred embodiment shown in a non-limiting manner in  FIG. 2 , the inner body  220  comprises an inner enclosure  222  comprising a permanent magnet  223  fixed to said inner enclosure  222  for rotation therewith. For example, said inner enclosure  222  comprises a spherical inner surface  224  that is closed and opposite the outer surface  221  of said inner enclosure  222 , and delimiting an inner volume. Said permanent magnet  223  has a magnetic moment that generally lies in an interval [0.1 A·m 2 , 1.5 A·m 2 ] (where “A·m 2 ” used herein denotes Amperes multiplied by square metres), has a substantially cylindrical shape, and is maintained, by bonding, welding or any other means, fixed to said inner surface  224  of the inner enclosure  222 . For example, the permanent magnet  223  is made of a Neodymium-iron-boron alloy grade N52 and is present in the shape of a cylinder measuring 8 cm in length with a radius of 3.36 cm. Such a configuration produces a magnetic moment of 1.5 A·m 2 . 
     Thus, when the satellite  100  is in orbit, said permanent magnet  223  tends to align itself with the Earth&#39;s magnetic field such that the inner enclosure  222  always substantially orients the same part of the outer surface  221  thereof towards the Earth&#39;s magnetic field. However, according to other examples not described herein, the permanent magnet  223  can also take on other forms, as can the inner surface  224  of the inner enclosure  222 . 
     In one more specific embodiment, said inner enclosure  222  comprises a substrate inside a volume that is not occupied by the permanent magnet  223 . 
     The use of a low-density substrate, for example made of polyurethane foam, is advantageous to ensure that the volume density of the inner body  220  reaches a predetermined value when, for example, the respective dimensions and materials of said inner enclosure  222  and of said permanent magnet  223  are prescribed without any possibility of modification during the manufacture thereof. For example, said polyurethane foam has a density that lies in an interval [240 kg·m −3 , 880 kg·m −3 ] (where “kg·m −3 ” used herein denotes kilograms per square metre). 
     Alternatively, in an alternative embodiment, the inner body  220  is solid. For example, the entire inner body  220  is a spherical permanent magnet. 
     It should be noted that the choice of materials used in the composition of the passive damper  200  is important. In this respect, these materials must be chosen such that they meet weight restrictions inherent to the astronautics field, as well as, in the case of the present disclosure, such that those forming the outer enclosure  210  are not magnetic. Furthermore, the materials of the passive damper  200  must meet robustness requirements, given that the inner body  220  is capable of coming into contact with the outer enclosure  210 . 
     Thus, in a preferred embodiment, at least one of either the inner surface  211  of the outer enclosure  210  and the outer surface  221  of the inner body  220  is made using a non-metallic material. Such a configuration is advantageous since it reduces the risks of attachment between the outer enclosure  210  and the inner body  220  in the event that the latter move closer to one another, for example as the result of electrostatic interactions caused by Van de Waals forces exerted between the respective atoms of said inner surface  211  and said outer surface  221 . 
     Preferably, the inner surface  211  of the outer enclosure  210  is made of a synthetic resin of the polyimide type. More particularly, said inner surface  211  is advantageously made of Vespel® which is a material certified for use in space. 
     It is also understood that, in order to limit contact between the inner body  220  and the outer enclosure  210 , said inner body  220  advantageously remains in suspension and centred in the viscous fluid  230 . 
     For this purpose, in a preferred embodiment, the densities of respectively the inner body  220  and of the viscous fluid  230  are substantially equal, at least for a temperature of said viscous fluid  230  within the interval [10° C., 30° C.] (where “° C.” used herein denotes degrees Celsius). In other words, for this temperature, for example 20° C., the buoyancy of the inner body  220  is substantially zero. 
     One additional advantage provided by such zero buoyancy of the inner body  220  within such a temperature interval is the ability to perform operational tests on the passive damper  200  when the latter is placed in a gravitational field, for example 1 g (or, in an equivalent manner, 9.8 m·s −2 ) at the Earth&#39;s surface, which is less restrictive, in particular from an implementation perspective, than operational tests in a microgravity environment. 
     In one specific embodiment of the invention, the pressure of the viscous fluid  230  is substantially equal to 5 bar for at least one temperature that lies in the interval [10° C., 30° C.]. Such a positive pressure in such a temperature interval is advantageous since it prevents any vaporisation of said viscous fluid  230  when the satellite  100  is:
         either in operation in an operational orbit, which corresponds to temperatures of the fluid  230  that lie in an interval [0° C., 50° C.],   or in a spent configuration orbit, which corresponds to temperatures of the fluid  230  that lie in an interval [−75° C., 0° C.], the inventors in particular having observed that said pressure remains substantially positive and equal to 1 bar for at least one temperature that lies in an interval [−75° C., −65° C.].       

     Preferably, the viscous fluid  230  is silicon oil of the dimethicone type. More particularly, said viscous fluid  230  is advantageously of the type PSF-5 cSt®, where the variations in density, dynamic viscosity, and kinematic viscosity of this fluid  230  in an interval [−75° C., 50° C.] are known to a person skilled in the art. However, according to other examples not described herein, other viscous fluids  230  can also be considered. 
     In the description hereafter and in a non-limiting manner, a situation is described wherein the satellite  100  is in orbit, rotating about an arbitrary axis of rotation, and comprises a passive damper  200  as shown in  FIG. 2 , the viscous fluid  230  being silicon oil of the dimethicone type. Moreover, said satellite  100  is close enough to the Earth for the inner magnet  223  to cooperate with the Earth&#39;s magnetic field according to the same principle as a compass. Thus, said inner body  220  is subjected to a magnetic torque that is equal to the vector product of the magnetic moment of the inner magnet  223  by the intensity of the Earth&#39;s magnetic field. 
     Since it is fixed to the satellite  100  for rotation therewith, the outer enclosure  210  is thus animated by a rotational movement that is identical to that of the satellite  100 , said rotational movement being firstly transmitted to the viscous fluid  230  in contact with the inner surface  211  of said outer enclosure  210 , then secondly to the inner body  220 , the outer surface  221  whereof is also in contact with said viscous fluid  230 . 
     Thus, said inner body  220  is not driven by the rotational movement of the viscous fluid  230  since it is, on the one hand, held back by the permanent magnet  223  which tends to remain aligned with the Earth&#39;s magnetic field, and on the other hand, substantially centred in said fluid  230 , and thus not in contact with the outer enclosure  210  as a result of the zero buoyancy thereof. Such a configuration of said passive damper  200  is particularly advantageous since, as long as the outer enclosure  210  is rotating, the fluid  230  is subjected to shear stresses originating from the differential in rotational speed between said outer enclosure  210  and said inner body  220 , these shear stresses ensuring the dissipation of energy associated with the rotational movement of the outer enclosure  210 , and thus of the satellite  100 . 
     It should be noted that, in the specific case in which the outer enclosure  210  and the inner body  220  are two concentric spheres respectively having the radii R 1  and R 2 , and for respective rotational speeds Ω 1  and Ω 2 , and in which the viscosity of the fluid is denoted μ, the viscous torque exerted on the inner body  220  can be estimated using the following formula: 
     
       
         
           
             T 
             = 
             
               
                 - 
                 8 
               
               ⁢ 
               πμ 
               ⁢ 
               
                 
                   
                     R 
                     1 
                     3 
                   
                   ⁢ 
                   
                     R 
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     It is furthermore understood that the more the rotational speed of the outer enclosure  210  increases, the higher the shear stresses, and the more the damping torque increases, which advantageously allows the rotation of the satellite  100  to be slowed down. In this respect, the inventors have observed that when the damping torque became substantially equal, as an absolute value, to the magnetic torque exerted on the inner body  220 , said inner body  220  was, in turn, driven by the rotational movement of the fluid  230 . This results in a decrease in the effectiveness of the passive damper  200 , which should be limited by dimensioning the passive damper  200  accordingly, as described, for example, hereinbelow. 
     Finally, it should be noted that although the passive damper  200  is designed so as to preferably prevent any contact between the outer enclosure  210  and the inner body  220 , in particular by means of a substantially zero buoyancy of said inner body  220  for at least one temperature of said viscous fluid  230  falling in the interval [10° C., 30° C.], this technical characteristic can become invalid when the satellite  100  is in orbit. This is, for example, the case when said passive damper  200  is subjected to angular accelerations when it is far from the centre of mass of the satellite  100 , these angular accelerations being capable of amplifying the effect of a buoyancy differential, in particular when cold when the density of the viscous fluid  230  becomes greater than the density of the inner body  220 . 
     However, the inventors have observed, by digital simulations, that when the inner body  220  and the outer enclosure  210  are in contact and moving relative to one another, said passive damper  200  is still suitable for generating a damping torque. In particular, they have established that in the case wherein the outer enclosure  210  and the inner body  220  are two concentric spheres, the damping torque increases with the off-centring between the respective centres of said two spheres. Thus, it is understood that the passive damper  200  is effectively still suitable for slowing the rotation of the satellite  100 . 
     In the precise example embodiment of the passive damper  200  in  FIG. 2 , the outer enclosure  210  and the inner enclosure  222  are made of Vespel®. The outer and inner radii of the outer enclosure  210  are respectively equal to 5.25 cm and 5 cm. The outer and inner radii of the inner enclosure  222  are respectively equal to 4.9 cm and 4.65 cm. The permanent magnet  223  is cylindrical in shape, has a length of 8 cm and a radius of 3.26 mm, and has a magnetic moment that is equal to 1.5 A·m 2 . The viscous fluid  230  is dimethicone oil of the type PSF-5 cSt®, the passive damper  200  comprising a mass of 29 g thereof. Moreover, the inner enclosure  222  and the permanent magnet  223  have a cumulated mass of 458 g, and said inner enclosure  222  comprises polyurethane foam having a density of 684 kg·m −3 , in a volume not occupied by said permanent magnet  223  and in a sufficient quantity for the inner body  220  to have a density that is substantially equal to the density of the fluid  230  at 20° C. 
     The inventors have observed that such a configuration of the passive damper  200  ensures sufficient slowing of the rotation of a satellite  100 , the inertia whereof is substantially equal to 100 kg·m 2 , placed in a circular polar orbit at a height of 700 km, and the active attitude control means whereof have malfunctioned, in a time interval varying from 1 week to 6 months. Such a time interval is particularly advantageous since it involves, for example, deorbiting a satellite  100  at the end of its life from a satellite constellation. 
     It is thus understood that the dimensioning of the passive damper  200  as well as the materials used in the composition thereof and the type of viscous fluid  230  are chosen so as to ensure slowing of the rotational speed of the satellite  100 , preferably stopping said rotation by a fixed deadline (for example 6 months), but also to guarantee that said device  200  remains functional, even at low temperatures, for example up to −75° C., when all energy sources on-board the satellite  100  are spent. 
       FIGS. 3 and 4  diagrammatically show an alternative embodiment of a passive damper  200  in  FIG. 2 , wherein the inner body  220  is capable of moving in rotation inside the outer enclosure  210  about a single axis of rotation (shown by a dotted line in  FIGS. 3 and 4 ). 
     For example, and as shown in  FIGS. 3 and 4 , the inner surface  211  of the outer enclosure  210  and the outer surface  221  of the inner body  220  each have a single axis of revolution that is substantially aligned with said single axis of rotation, and are furthermore substantially proportional. 
     In such a configuration, the outer surface  221  of the inner body  220  also advantageously has, at the axis of rotation, protrusions  225 , the respective ends whereof are housed, in a contactless manner, inside cavities  215  made, at the axis of rotation, in the inner surface  211  of the outer enclosure  210 . For example, and as shown in a non-limiting manner in  FIG. 3 , the outer surface  221  of the inner body  220  comprises two protrusions  225 , the respective ends whereof are pointed, said protrusions being respectively positioned at the level of two points of intersection of the axis of rotation with said outer surface  221 . These protrusions  225  associated with said cavities  215  have the advantage, in addition to the choice of appropriate material for the inner surface  211  of the outer enclosure  210 , and the outer surface  221  of the inner body  220 , of minimising the risks of attachment caused by the Van der Waals forces between these two surfaces. However, according to other examples not described herein, the end of the protrusions can also be non-pointed. Moreover, such arrangements are not limited to the configuration in  FIG. 3 . 
     In the example shown in  FIG. 3 , according to a sectional plane that passes through the axis of rotation, the inner surface  211  of the outer enclosure  210  and the outer surface  221  of the inner body  220  are respectively cylindrical in shape. More particularly, the inner body  220  comprises an inner enclosure  222 , comprising a permanent magnet  223 , and the outer surface  212  of the outer enclosure  210  in addition to the inner surface  224  of the inner enclosure  222  are also respectively cylindrical in shape. However, according to other examples not described herein, the outer surface  212  of the outer enclosure  210  and the inner surface  224  of the inner enclosure  222  can also take on other shapes. 
     In the example shown in  FIG. 4 , according to a sectional plane that passes through said single axis of rotation, the inner surface  211  of the outer enclosure  210  and the outer surface  221  of the inner body  220  are respectively H-shaped. More particularly, the inner body  220  is solid and the outer surface  212  of the outer enclosure  210  is also H-shaped. Such a configuration in particular allows the shearing surface to be increased for identical overall dimensions. 
     It is understood that such an alternative embodiment allows the freedom of movement of the inner body  220  inside the outer enclosure  210  to be limited, such that the passive damper  200  is intended to slow the component of the rotational movement of a satellite  100  projecting on said single axis of rotation. 
       FIG. 5  shows a preferred alternative embodiment of the satellite  100 , wherein said satellite  100  comprises three passive dampers  200  as shown in  FIG. 3 , said three passive dampers  200  being positioned on the inner surface  112  of the main body  110 , and arranged such that the axes of rotation of the respective inner bodies  220  thereof are not parallel in pairs. Such a configuration is equivalent, from a theoretical perspective, to that of a passive damper  200  comprising at least one inner body  220  capable of moving in rotation inside the outer enclosure  210  about arbitrary axes of rotation. However, according to other examples not described herein, three passive dampers  200  different to those shown in  FIG. 3  can also be used, for example passive dampers  200  such as those shown in  FIG. 4 . More generally, three passive dampers  200  can be used, configured such that the respective inner bodies  220  thereof are capable of moving in rotation about a single axis of rotation, and arranged such that these axes of rotation are not parallel in pairs. 
     Moreover, the inventors have observed that the present invention could be implemented by means of two passive dampers  200 , configured such that the respective inner bodies  220  thereof are capable of moving in rotation about a single axis of rotation, and arranged such that these axes of rotation are not parallel in pairs. More specifically, such a configuration produces passive attitude control means, the performance whereof is substantially equal to that of a configuration comprising three passive dampers  200  of the same type. More specifically, and in practice, when the satellite  100  is in the spent configuration thereof, it is animated by a rotational movement, at a high speed, according to a plurality of axes of rotation. Even if the satellite  100  were animated by a rotational movement according to a single axis of rotation, the probability that this single axis of rotation is contained within a plane orthogonal to the plane defined by the axes of rotation of the two inner bodies  220  is negligible. 
     In addition, the passive attitude control means can also comprise a single passive damper  200  configured such that the inner body  220  thereof is capable of moving in rotation about a single axis of rotation. The inventors have observed that such a configuration would procure a high-performance reduction in the rotational speed of the satellite  100  in the spent configuration thereof. 
     Generally speaking, it should be noted that the embodiments considered above have been described by way of non-limiting examples, and other alternative embodiments can thus be envisaged.