CENTER OF INERTIA BALANCING SYSTEM FOR A TELECOMMUNICATIONS SATELLITE.

Abstract

A balancing device for the center of mass O of a telecommunications satellite placed in geostationary orbit in a local orbital frame (X, Y, Z) centered at O, where X is the axis parallel to a velocity vector of the satellite's movement, is proposed. The satellite comprises a body with two faces (1222, 1224) parallel and orthogonal to X, and a payload instrument fixed to face 1222 with a given mass and length. The center of mass is displaced by a perturbation acting on the satellite. The balancing device comprises: a deployment device fixed to face 1224, the deployment device having a length greater than the instrument's length; a balancing element fixed to the end of the deployment device, the balancing element having a mass less than the instrument's mass; and the deployment device being configured to move the balancing element along X. Figure for the abstract: Fig. 4

Description

Title of the invention: System for balancing the center of inertia of a telecommunications satellite technical field

[0001] The invention relates generally to telecommunications satellites and in particular to a system for balancing the center of inertia of a telecommunications satellite circulating in a geostationary orbit.

[0002] A telecommunications satellite typically comprises a platform and a payload. The payload of a telecommunications satellite includes various instruments such as transmitting antennas, receiving antennas, and electronic systems for receiving, processing, and transmitting the payload signal and associated data. The platform encompasses all systems dedicated to the satellite's operation, excluding systems and functionalities related to the payload signal. The platform includes a body, a chemical and / or electrical propulsion system, an attitude and orbit control system, and an electrical power generation and management system. The entire telecommunications satellite is configured to position the satellite's center of mass in its equilibrium position, substantially at the midpoint of the body.

[0003] The lifespan and success of a satellite mission are strongly linked to energy constraints, particularly the depletion of propellant (liquid propellants) and the reliability of the electrical power supply. The electrical power generation and management system maintains the platform and payload systems in operational condition, while the propulsion system places the satellite in geostationary orbit, maintains it on its nominal trajectory, and orients its instruments.

[0004] The satellite, orbiting in geostationary orbit, is linked to a local orbital frame R consisting of three conventional axes X, Y, and Z, centered on the satellite's center of mass O in its equilibrium position. The frame R = (X, Y, Z) is a right-handed orthogonal frame centered on the satellite's center of mass in its equilibrium position and in circular translation in the Earth's geocentric frame of reference. The X-axis corresponds to the axis parallel to the satellite's velocity vector in the geostationary orbit in the geocentric frame of reference. The Z-axis is directed towards the Earth and corresponds to the axis passing through the satellite's center of mass and the Earth's center of gravity. Furthermore, the Y-axis is orthogonal to the orbital plane, that is, to the X and Z axes.

[0005] The satellite's attitude and orbit control system is used to counteract the effects of disturbances acting on the satellite that may have consequences significant disturbances can affect the proper functioning of the satellite. Such disturbances can have external origins, meaning they are caused by phenomena external to the satellite, such as aerodynamic, magnetic, radiative, or gravitational disturbances. Alternatively, these disturbances can have internal origins when they are linked to mechanisms or displacements inherent to the satellite, such as the movement of mechanical parts or a decrease in propellant for the thrusters. For example, a disturbance acting on the satellite can cause a displacement of the satellite's center of mass, denoted O', relative to the center O of the local orbital frame R=(X, Y, Z), the satellite's center of mass in its equilibrium position.

[0006] Attitude control includes all processes related to controlling the orientation of the satellite in the local orbital frame R=(X, Y, Z). Satellite orbit control includes all processes associated with positioning the satellite and maintaining its trajectory in its geostationary orbit.

[0007] Conventionally, each of the orbital parameters of the geostationary orbit can be controlled by performing one or more orbital maneuvers using the propulsion system in order to adjust the satellite's displacement in its geostationary orbit according to perturbations. Such control notably takes into account the longitude of the satellite's orbit.

[0008] In order to control the orbit of the satellite, the displaced center of inertia O' of the satellite with coordinates (XO', YO', Zo ) in the local orbital frame R with center O (Xo, Yo, Zo ), must be included in a control volume defined by parameters (Xmax, Ymax, Zmax).

[0009] Typically, after the satellite is positioned in its geostationary orbit following launch, the payload instruments are deployed. The deployment of the instruments produces an internal disturbance that triggers a maneuver plan involving the propulsion system in order to maintain the satellite's orbit.

[0010] However, the deployment of a payload instrument, such as a telecommunications antenna including a large diameter reflector, along a given axis direction, for example the X axis, generates a significant asymmetry of the satellite on this axis with respect to the center of inertia O.

[0011] The asymmetry of the satellite about the given axis (X-axis, for example) with respect to the center O of the local orbital frame R causes a displacement of the center of inertia O' of the satellite outside the required control volume, such that IX0 - Xol is greater than IXmaxl or IX0 - Xol = IXOI > IXmaxL. Furthermore, the control of the satellite's orbit is primarily defined by the control of the satellite's longitude in geostationary orbit in this case. Outside the required control volume, the satellite's orbit is controlled by means of the propulsion system and by adjusting the thrust direction component tan- gentle to the orbit, is then impossible.

[0012] The known solutions for adjusting the displaced center of inertia O' out of the control volume required to address asymmetry compensation rely on major technical modifications of the propulsion device of telecommunications satellites.

[0013] For example, it is known to add thrusters to the fixed-engine propulsion system. However, such a solution significantly increases the amount of propellant to be added to the satellite.

[0014] In addition, these solutions can be associated with payload rearrangements via a rotation of the satellite axes, by 90° for example, and / or propulsion in order to symmetrize the satellite after deployment.

[0015] Document FR1455630 Al proposes, for example, a very important modification of the propulsion device in steerable motors allowing the translation and / or rotation of the propulsion via multi-axis arms in the local orbital frame R.

[0016] Such solutions based on modifications to the propulsion system all entail a high cost related to the modifications of the propulsion system or the satellite, as well as a reduction in the theoretical lifespan of the satellite.

[0017] Indeed, the propulsion device can only function correctly to control the satellite's orbit if the satellite's center of inertia is within the required control volume, constrained by the arrangement of the thrusters on the satellite.

[0018] There is therefore a need for a system capable of improving the orbit control of a telecommunications satellite. Summary of the invention

[0019] The present invention improves the situation by proposing a device for balancing the center of inertia of a telecommunications satellite placed in geostationary orbit in a local orbital frame having as its center O the center of inertia of the telecommunications satellite in its equilibrium position, and comprising three axes X, Y and Z, the X axis being parallel to a velocity vector of the satellite's movement in the geostationary orbit, the Z axis being directed towards the Earth and the Y axis being orthogonal to the orbital plane (X, Z), the telecommunications satellite comprising at least one body having a shape comprising at least two parallel faces substantially orthogonal to the X axis, and a payload instrument fixed to the telecommunications satellite on one of the two faces of the body substantially orthogonal to the X axis, called the instrument attachment face,The payload instrument, having a given mass and length (LiS), has its center of mass displaced to a point O' by a perturbation acting on the telecommunications satellite. The balancing device further includes:

[0020] - a deployment device attached to the telecommunications satellite by the face of the body substantially orthogonal to the X axis and opposite the instrument's fixing face, the deployment device having a given length L^ greater than the instrument length Lj,

[0021] - a balancing element fixed to the end of the deployment device, the element balancing having a given mass less than the mass of the instrument M(.

[0022] The deployment device is configured to move the balancing element in at least one direction parallel to the X axis.

[0023] In one embodiment, the balancing element may comprise a metallic material.

[0024] The balancing element may have a compact shape.

[0025] In particular, the compact form of the balancing element can be a rectangular prism.

[0026] Alternatively, the compact form of the balancing element can be a cylinder.

[0027] The mass of the balancing element can be equal to the mass of the instrument M divided by two, the length of the deployment device Ld being equal to the length of the instrument L, multiplied by two.

[0028] In some embodiments, the balancing element may include at least one other payload instrument.

[0029] The deployment device may include a telescopic arm.

[0030] In a particular embodiment, the telescopic arm may comprise at least two sections connected by a sliding link having one degree of freedom, the deployment device moving the balancing element by performing a translation in a direction parallel to the X axis.

[0031] Alternatively, the deployment device may include a multi-axis and multi-arm mechanism.

[0032] In a particular embodiment, the multi-axis and multi-arm mechanism may comprise at least two sections connected to each other and to the body by pivot links, each having one degree of rotational freedom, the deployment device moving the balancing element in at least the cutting plane (X, Y).

[0033] The telecommunications satellite may include a plasma propulsion device.

[0034] In some embodiments, the telecommunications satellite may include a solar power generation system and a balancing device control unit, the balancing device being actuated by command from the control unit using energy supplied by the solar power generation system.

[0035] The invention also provides a method for actuation of the center of inertia balancing device of a telecommunications satellite comprising the steps consisting of:

[0036] - to launch the satellite and position the satellite in geostationary orbit

[0037] - deploy the payload instrument attached to the telecommunications satellite by one of the two faces of the body substantially orthogonal to the X axis;

[0038] - deploy the balancing device attached to the telecommunications satellite by the face of the body substantially orthogonal to the X axis opposite the fixation face of the instrument.

[0039] The method further comprises at least one iteration of the following steps, in response to a disturbance or a maintenance operation of the communication satellite:

[0040] - apply calibration maneuvers to the telecommunications satellite and calculate the displacement of the center of inertia O' relative to its equilibrium position O;

[0041] - apply a process to define a maneuver plan for the device balancing correction of the displaced center of inertia O';

[0042] - actuate the center correction balancing device maneuvers of inertia.

[0043] In one embodiment, the maneuver plan can be determined by a ground station and transmitted to the telecommunications satellite.

[0044] In particular, the maneuver plan can be implemented by the telecommunications satellite comprising an autonomous processing and control unit.

[0045] The method and device for balancing the center of inertia of a telecommunications satellite according to the embodiments of the invention make it possible to remedy all or part of the limitations of the solutions of the prior art, in particular those set out above, by proposing a solution which makes it possible to compensate for asymmetry of the satellite following the deployment of an instrument of the payload, in particular by avoiding modifications of the propulsion device and by increasing the theoretical lifetime of the satellite. Description of the figures

[0046] Other features, details and advantages of the invention will become apparent from the description made with reference to the accompanying drawings given by way of example.

[0047] [Fig.1] Fig.1 illustrates a telecommunications satellite circulating in a geostationary orbit around the Earth.

[0048] [Fig.2] Fig.2 is a diagram representing the telecommunications satellite and its components illustrated in [Fig.1].

[0049] [Fig.3] Fig.3 is a simplified diagram representing a balancing system in the telecommunications satellite, according to embodiments of the invention.

[0050] [Fig.4] Fig.4 illustrates two section planes on the (X, Y) and (X, Z) planes of the telecommunications satellite with a receiving antenna and a balancing device deployed.

[0051] [Fig. 5] [Fig. 5] illustrates the body and balancing device of a television satellite communications, according to an embodiment of a "telescopic" deployment device.

[0052] [Fig. 6] [Fig. 6] illustrates the body and balancing device of a television satellite communications, according to an embodiment of a "multi-axis and multi-arm" deployment device.

[0053] [Fig.7] Fig.7 is a diagram representing the steps of an actuation process of the balancing system for controlling the position of the center of inertia of a telecommunications satellite.

[0054] Identical reference numerals are used in the figures to designate identical or analogous elements. For clarity, the elements shown are not to scale. Detailed description

[0055] Fig. 1 schematically represents a telecommunications satellite 10 circulating in a geostationary orbit 20 around the Earth 30, defined by definition at an altitude of 35,786 km.

[0056] Satellite 10 is configured to perform a primary telecommunications mission.

[0057] The positioning of the satellite 10 is defined by means of a local orbital frame R=(X, Y, Z) centered at the center of inertia O of the satellite 10 when it is in its equilibrium position.

[0058] The local orbital frame R has three axes X, Y and Z. The X axis is parallel to a velocity vector corresponding to the velocity vector of the satellite 10 moving in the geostationary orbit 20. The Z axis is directed towards the Earth 30 and the Y axis is orthogonal to the orbital plane (X, Z).

[0059] As shown in [Fig. 2], the satellite 10 comprises a platform 120 and a payload 140. The payload 140 is connected to the platform 120 and includes all the systems and equipment intended to be used by the satellite 10 to perform its mission. The payload 140 includes at least one instrument 142. This instrument 142 may be, for example, but is not limited to, a telecommunications antenna, including, for example, a large-diameter reflector (e.g., a 9 m reflector). The platform 120 of the satellite 10 is the service module that provides power to the satellite 10 and enables the satellite 10 to be positioned and maintained in geostationary orbit 20. The platform 120 includes at less a body 122, an electrical energy production and management system 124 and a propulsion device 126 (not shown in [Fig.l]).

[0060] The body 122 of the satellite 10 has a shape comprising at least two faces 1222, 1224 parallel to each other. As shown in [Fig. 1], without limitation, the body 122 of the satellite 10 may have the shape of a rectangular parallelepiped with six faces that are parallel in pairs. For example, the two parallel faces 1222, 1224 may be substantially orthogonal to the X-axis.

[0061] The payload 140 of the satellite 10 may include an instrument 142 attached to the satellite 10 on one of the two parallel faces 1222, 1224 of the body 122, which are substantially orthogonal to the X-axis. As shown in Figures 1 and 2, the instrument 142 is attached to the satellite 10 by the face 1222, called the instrument 1222 attachment face. The instrument 142 has a given instrument mass and a given instrument length Lz. The instrument 142 may further include a position of the center of mass (X, Y) in the cross-sectional plane (X, Y) of the local orbital frame R.

[0062] The center of mass O of the satellite 10 in its equilibrium position with coordinates (Xo, Yo, Zo) is then positioned such that the points Xo and Yo are located at the midpoint of the rectangular parallelepiped body 122, in the cutting plane (X, Y). Furthermore, the point Zo of the center of mass O is positioned approximately on the axis of symmetry of the body 122, parallel to the Z axis.

[0063] The electrical energy production and management system 124 of the platform 120 may include, without limitation, solar panels as illustrated in [Fig. 1] for producing electrical energy by the photovoltaic effect. Two solar panels may, for example, be used. These two panels may be fixed directly to the body 122 of the platform 120, respectively on two opposite faces of the body 122. For example, a first solar panel may be arranged on the "+Y face" while a second solar panel may be fixed on the opposite "-Y face", substantially orthogonal to the Y axis.

[0064] The "+Z face" corresponds to the face of the rectangular parallelepiped body 122 directed towards the Earth 30, parallel to the "-Z face" arranged on the side opposite the Earth 30, the +Z and -Z faces being substantially orthogonal to the Z axis. For example, and in no way limiting, the payload 140 may include a second telecommunications antenna which may be carried by the "+Z face" (the second telecommunications antenna is not shown in the figures).

[0065] The two other opposite faces of the rectangular parallelepiped body 122 are designated by the "+X face" and the "-X face" and are substantially orthogonal to the X axis. In the example shown in [Fig. 1], the mounting face of the instrument 1222, which carries the instrument 142 of the payload 140, is the "+X face" of the satellite 10. Similarly, instrument 142 can be carried by the "-X face" of satellite 10 (such an example is not shown in the figures).

[0066] As used here, the center of inertia coincides with the center of mass, and with the center of gravity in this configuration of satellite 10, as illustrated by [Fig.1], at this given altitude of satellite 10 relative to the Earth 30.

[0067] The satellite 10 is initially launched into space by a launcher until it reaches its geostationary orbit 20. The satellite 10 is said to be launched. The launcher includes, in particular, a fairing located at its forward end, with an aerodynamic profile. The fairing's purpose is, among other things, to protect the satellite 10 from inclement weather and extreme temperature or humidity conditions, both on the ground and at the start of the launch. After launch, the fairing is jettisoned as soon as the air density is sufficiently low so that the resulting aerodynamic forces cannot damage the satellite 10.

[0068] After the satellite 10 is positioned in its geostationary orbit 20 following its launch, the instrument 142 of the payload 140 is deployed. The deployment of an instrument 142, such as the telecommunications antenna comprising a large-diameter reflector (Figure 1), can generate a strong asymmetry of the satellite 10 along the X-axis with respect to the center O of the local orbital frame R.

[0069] In this configuration of satellite 10, as shown in [Fig. 1], such an asymmetry related to the deployment of instrument 142 of satellite 10 causes a displacement of the center of inertia O' of satellite 10 outside the required control volume, such as, for example, IXod > IXmaxL

[0070] Thus, the center of inertia O of the satellite 10 is displaced at the level of the point O' by a disturbance acting on the satellite 10.

[0071] In the satellite configuration 10 shown in Figure 1, the required control volume can in practice be reduced to a surface projected onto the (X, Y) plane orthogonal to the Z-axis of the local orbital frame R centered at O, also called the control surface. This control surface can, for example, have the shape of a rectangle with dimensions dX and dY, respectively along the X and Y axes. The dimensions of the body 122 along the X and Y axes, respectively, are denoted DX and DY. Thus, the characteristic threshold values ​​IXmaxl and IYmaxl can, for example, be defined according to the following equations (1) and (2):

[0072] 2x|XmJ=dX=(9%xDX) (1)

[0073] 2 x |Y^J = dY= (15% x DY) (2).

[0074] Thus, for an instrument 142 composed, for example, of a 9m reflector, the The placement of the center of inertia O' can be characterized by a value such that:

[0075] |XO. |=23%xffi(3)

[0076] Thus, the value |XO. | can be approximately equal to two and a half times the threshold value of IXmaxl required to maintain control of the longitude of the satellite 10 in geostationary orbit 20 via the propulsion device 126.

[0077] Embodiments of the invention provide a device and a method for balancing the displaced center of mass O' of the satellite 10 in the local orbital frame R centered at O, making it possible to maintain acceptable and effective conditions for using the propulsion device 126 as a means of orbit control for the satellite 10, and in particular longitude control. The embodiments of the displaced center of mass O' are located within a control volume defined by (Xmax, Ymax, Zmax). The balancing device 128 according to the invention is advantageously based on the displacement of a balancing element 1284 by a deployment device 1282, along at least one direction parallel to the X-axis, as for example in the case illustrated by [Fig. 1], in order to compensate for the displacement IXod of the center of mass O'.

[0078] Figure 3 is a simplified representation of the satellite 10 comprising a platform 120 and a payload 140. Figure 3 represents in particular the balancing device 128 of the center of inertia of the satellite 10, when the satellite 10 is placed in geostationary orbit 20. The positioning of the satellite 10 is defined in the local orbital frame R=(X, Y, Z) having center O (center of inertia of the satellite 10 in its equilibrium position).

[0079] Advantageously, the balancing device 128 comprises a deployment device 1282 having a first end fixed to the satellite 10 by one of the two parallel faces 1222 and 1224 of the body 122, substantially orthogonal to the X-axis. As shown in Figures 3, the balancing device 128 is fixed to the satellite 10 by the face 1224, called the attachment face of the device 1224, opposite the attachment face of the instrument 1222. The deployment device 1282 comprises a given mass and a given length Lj. The length of the deployment device Ld is defined such that Lrf is greater than the length of the instrument L;.

[0080] The balancing device 128 further includes a balancing element 1284 fixed at the second end of the deployment device 1282. The balancing element 1284 comprises a given mass and a given length L. The mass of the balancing element Me is defined such that Me is less than the mass of the instrument M.

[0081] The balancing device 128 therefore comprises a resultant mass + and a resultant length ( ) • In one embodiment, the resultant mass ( ) can be between 30kg and 150kg, and the resulting length (L^ + L^) can be between 5m and 20m. The balancing device 128 further includes a position of the center of mass (Xe, Ye) in the cutting plane (X, Y) of the local orbital frame R.

[0082] According to some embodiments, the length of the balancing element Lf can be considered substantially negligible compared to the length of the deployment device Lj, ​​while the mass of the balancing element can be equal to N times the mass of the deployment device Md, the multiple N being between three and ten, depending on the technology chosen to constitute the deployment device (Me = TV x Md).

[0083] The mass and length parameters of the balancing device 128 can be determined according to an iterative process, for example, from moment equilibrium calculations, considering in particular the position of the center of mass of the instrument (Xi, Yi) and the position of the center of mass of the balancing device (Xe, Ye), as well as their respective masses Mi and (Me -I-Mi)

[0084] In some embodiments, the mass of the balancing element may be equal to the mass of the instrument My divided by two), the length of the deployment device Lf / being equal to the length of the instrument Lz multiplied by two ( Lj = Ls X 2 ).

[0085] The deployment device 1282 is configured to move the balancing element 1284 in at least one direction parallel to the X axis.

[0086] The balancing device 128 makes it possible to compensate for the displacement of the center of inertia O' by a deployed instrument 142 generating the asymmetry of the satellite 10, such as the telecommunications antenna comprising a large diameter reflector.

[0087] The satellite 10 may include a control unit 1286 configured to control the movement of the balancing device 128 and actuate it by means of a control using the energy supplied by the solar power generation system 124. Thus, the balancing device 128 can operate without using the limited fuel resources of the satellite 10, thereby increasing the lifespan of the satellite 10 compared to prior art devices.

[0088] Advantageously, the balancing device 128 has by itself the capacity to modify the inertias of the satellite 10 to optimize attitude control and thus allow an active recentering of the displaced center of inertia O' of the satellite 10 in geostationary orbit 20 in order to optimize the consumption of propellants (for example xenon).

[0089] In embodiments, the balancing element 1284 may be made of a metallic material, which may be, for example and without limitation, steel, titanium or aluminum.

[0090] According to some embodiments, the balancing element 1284 may include or be composed of one or more other payload instruments (for example, a star sounder). In this case, the balancing element 1284 is referred to as 'active mass' as opposed to a 'dead mass' without instrument implementation.

[0091] The balancing element 1284 may have a compact shape, such as, for example, a cylindrical or rectangular prism shape. The shape of the balancing element 1284 may, for example, be chosen according to the fastening characteristics of the deployment device 1282.

[0092] In embodiments, the balancing element 1284 may have a shape adapted to the available volume under the fairing of the launcher of the satellite 10, which allows the launch of the satellite 10 into geostationary orbit 20.

[0093] In embodiments, the shape of the balancing element 1284 can further be chosen so as to modify the effects and the solar wind exposure of the satellite 10 when it is in geostationary orbit 20, which makes it possible to limit the variations of onboard angular momentum for attitude control, and thus to optimize pointing.

[0094] Figure 4 shows two cross-sectional views of the satellite 10 with the instrument 142 and the balancing device 128 deployed. The upper representation of Figure 4 shows the satellite 10 in the (X, Z) plane and the lower representation of Figure 4 shows the satellite 10 in the (X, Y) plane, where the instrument 142 is a telecommunications antenna comprising a large-diameter reflector directed towards the Earth 30.

[0095] The deployment device 1282 is thus fixed to the satellite 10 by the fixing face of the device 1224 substantially orthogonal to the X axis and opposite to the fixing face of the instrument 1222. Thus, if the instrument 142 is carried by the "+X face", the balancing device 128 will be carried by the "-X face", and conversely if the instrument 142 is carried by the "-X face", the balancing device 128 will be carried by the "+X face".

[0096] According to one embodiment, the balancing device 128 of the satellite 10 may include a so-called "telescopic" deployment device 1282, also called a telescopic arm, as illustrated in [Fig. 5]. In this embodiment, the telescopic arm 1282 comprises at least two sections 1282-1, 1282-2. The sections 1282-1, 1282-2 have a cross-section, and this cross-section is, for example, circular or square. The sections 1282-1, 1282-2 are connected by a so-called telescopic link 1282-3 having one degree of freedom. The balancing element 1284 is then moved by the deployment device 1282 by implementing a translational movement, along a direction parallel to the X-axis. In a non-limiting example, the telescopic arm 1282 may be characterized by a mechanism using a compressed carbon lattice.In general, the mechanism for implementing the translational movement is configured to allow active movement of the balancing element 1284, i.e., adjustable after deployment.

[0097] According to certain embodiments, a first section 1282-1 of the telescopic arm 1282 is connected to the body 122 of the satellite 10 by the mounting face of the device 1224, which is substantially orthogonal to the X-axis. This first section 1282-1 is stationary relative to the body 122 of the satellite 10. A second section 1282-2 of the telescopic arm 1282 is connected to the balancing element 1284 and is stationary relative to the balancing element 1284. The sliding joint 1282-3 allows the second section 1282-2 to retract into the first section 1282-1 and the body 122. Advantageously, the stored volume of the balancing system is then very compact since it is limited to the balancing element 1284 and the first section 1282-1 of the telescopic arm 1282.

[0098] In another embodiment, the balancing device 128 of the satellite 10 may include a deployment device 1282 referred to as a "multi-axis and multi-arm" 1282, also called a multi-axis and multi-arm arm, as illustrated in [Fig. 6]. In this embodiment, the multi-axis and multi-arm arm is a mechanism derived from telecommunications antenna deployment mechanisms comprising a large-diameter reflector.

[0099] The multi-axis, multi-arm 1282 may comprise several sections (comprising at least two sections), connected to each other and to the body 122 of the satellite 10 by pivot joints, each having one degree of rotational freedom. For example, in [Fig. 6], three sections 1282-4, 1282-5, 1282-6 are connected to each other by three pivot joints 1282-7, 1282-8, 1282-9, which may be characterized by mechanisms with electric motors and / or springs. The balancing element 1284 can then be moved by the deployment device 1282, and depending on the choice of pivot joint technology, this movement may be either active (i.e., adjustable after deployment) or passive (i.e., not modifiable after deployment).

[0100] According to certain embodiments, the axis of rotation of the pivot joints 1282-7, 1282-8, 1282-9 of the multi-axis, multi-arm arm 1282 is substantially parallel to the Z-axis when the satellite 10 is in geostationary orbit 20 according to the mission attitude. A first section 1282-4 of the multi-axis, multi-arm arm 1282 is connected to the body 122 of the satellite 10 on the mounting face of the device 1224 by the joint 1282-7. This first section 1282-4 is movable relative to the body 122 in the (X, Y) plane. A second section 1282-5 of the multi-axis, multi-arm arm 1282 is connected to the first section 1282-4 of the multi-axis, multi-arm arm 1282 by the joint 1282-8. A third section 1282-6 of the multi-axis and multi-arm 1282 is connected to the second section 1282-5 of the multi-axis and multi-arm 1282 by the link 1282-9.The third section 1282-6 of the multi-axis and multi-arm 1282 is connected to the balancing element 1284, and is stationary relative to the balancing element. 1284. The pivot links 1282-7, 1282-8, 1282-9 allow the sections 1282-4, 1282-5, 1282-6 and the balancing element 1284 to move in the (X, Y) plane.

[0101] According to other embodiments, the axis of rotation of the pivot joints 1282-7, 1282-8, 1282-9 of the multi-axis and multi-arm 1282 can be further provided with additional components along the X-axis and / or the Y-axis. In this case, the pivot joints 1282-7, 1282-8, 1282-9 are configured to allow the displacement of sections 1282-4, 1282-5, 1282-6 in the local orbital frame R=(X, Y, Z) so as to perform a finer adjustment of the balancing element 1284. Such an embodiment of the pivot joints is particularly useful when the balancing element 1284 is an active mass.

[0102] Thus, during the equilibrium calculations for determining the characteristic mass and length parameters of the balancing device 128, cost and size factors (available storage volume on the satellite 10 at launch) can be taken into account. For example, since the size of the deployment device 1282 using telescopic arm technology is smaller than the size of the multi-axis and multi-arm technology, the length of the telescopic arm can be greater than that of the multi-axis and multi-arm. Consequently, the mass of the balancing device 128, and in particular of the balancing element 1284, will be lower with this technology. It should be noted that the mass of the telescopic arm can also be lower than that of the multi-axis and multi-arm for the same length.

[0103] The embodiments of the invention thus provide an efficient and inexpensive balancing device 128, not requiring major modifications to the platform 120 and in particular no modification to the propulsion device 126. Attaching the balancing device 128 to the body 122 of the satellite 10 limits the overall size of the satellite 10, while its use has a limited impact on the fuel consumption of the satellite 10. Advantageously, the embodiments of the invention are not intrusive on the satellite 10.

[0104] According to some embodiments, the satellite 10 comprises a plasma propulsion system 126. As used here, plasma propulsion refers to a type of plasma-based space propulsion system. This type of propulsion system uses variable electromagnetic fields and radiation to heat, ionize, and accelerate a gas, such as, for example, but not limited to, xenon, krypton, or iodine. Plasma propulsion allows for the use of a lighter or more powerful satellite 10, by reducing the mass of onboard propellant.

[0105] The [Fig.7] is a flowchart representing a method of actuation of the balancing device 128 implemented to control the position of the center of inertia of a satellite 10.

[0106] The initialization phase of the balancing device 128 of the satellite 10 in geostationary orbit 20 may include a step 720 in which the satellite 10 is launched and positioned in geostationary orbit 20.

[0107] The implementation of the balancing device 128 and therefore the use of a propulsion device 126 having a plasma propulsion has an impact of only a few days on the time of positioning (or putting into orbit) of the satellite 10 compared to the prior art systems, due in particular to the realization of the deployment of the balancing device 128 during the test phases and after positioning.

[0108] All devices deployed after the satellite 10 is placed into orbit are called "space appendages." Space appendages may include solar panels, telecommunications antennas or reflectors, the balancing device 128, or any other equivalent equipment of the satellite 10. During the launch and placement of the satellite 10 into orbit, the various space appendages are subjected to significant stresses. To protect them during this phase, the space appendages can be folded and temporarily held against at least one face of the satellite 10's body 122 to immobilize them. The appendages can be "stacking," that is, held in a storage configuration, by a stacking device comprising several holding and release mechanisms. After this phase, the stacking device is actuated to release the space appendages and ensure their deployment.

[0109] The initialization phase of the balancing device 128 of the satellite 10 in geostationary orbit 20 may further include a step 740 of deployment of an instrument 142 fixed to the satellite 10 by one of the two faces 1222, 1224 of the body 122 substantially orthogonal to the X axis. For example, a telecommunications antenna, including a large diameter reflector, may be fixed to the fixing face of the instrument 1222.

[0110] The initialization phase of the balancing device 128 of the satellite 10 in geostationary orbit 20 can also include a step 760 of deployment of the balancing device 128 fixed to the satellite 10 by the fixing face of the device 1224 substantially orthogonal to the X axis, opposite to the fixing face of the instrument 1222.

[0111] After the initialization phase, steps can be iterated to respond to a disturbance or to perform a maintenance operation on satellite 10.

[0112] Such steps may include:

[0113] - a calibration step 782, in which the satellite 10 performs maneuvers of calibration of satellite 10 in order to calculate the displacement of the center of inertia O' relative to its equilibrium position O.

[0114] - a step 784 of processing information from the calibration step in order to define a maneuver plan for the balancing device 128 to correct the displaced center of inertia O'.

[0115] According to some embodiments, the maneuver plan is implemented by satellite 10, which includes an autonomous processing and control unit. The maneuver plan can also be determined by a ground station and transmitted to satellite 10.

[0116] The steps for responding to a disturbance or a maintenance operation 80 of the satellite 10 may further include a step 786 of actuation of the maneuvers of the balancing device 128 for correcting the center of inertia.

[0117] The invention is not limited to the embodiments described above by way of non-limiting example. It encompasses all variant embodiments that can be envisaged by a person skilled in the art. In particular, a person skilled in the art will understand that the invention is not limited to the types of deployment devices 1282 described or to the forms of balancing element 1284 described, by way of non-limiting example.

Claims

Demands

1. A satellite telecommunications system (10) comprising a balancing device (128) for the center of mass of said satellite system (10) placed in geostationary orbit (20) in a local orbital frame centered at O, the center of mass of the satellite system (10) in its equilibrium position, said frame having three axes X, Y, and Z, the X axis being parallel to a velocity vector of the satellite system (10) in the geostationary orbit (20), the Z axis being directed towards the Earth (30), and the Y axis being orthogonal to the orbital plane (X, Z), said satellite system (10) further comprising a body (122) having a shape comprising at least two parallel faces (1222, 1224) substantially orthogonal to the X axis, and a payload instrument (142) attached to the satellite system (10) on one of the two substantially orthogonal faces of the body. to the X axis, called the instrument mounting face (1222),said payload instrument (142) having a given mass M, and a given length Lz, the center of inertia of the satellite system (10) being displaced to a point O' by a perturbation acting on the satellite system (10), characterized in that the balancing device (128) comprises:, - a deployment device (1282) attached to the satellite system (10) by the face (1224) of the body substantially orthogonal to the X axis and opposite said attachment face of the instrument (1222), said deployment device (1282) having a given length Lt / greater than the length of the instrument U - a balancing element (1284) fixed to the end of said deployment device (1282), said balancing element (1284) having a given mass less than the mass of the instrument said deployment device (1282) being configured to move said balancing element (1284) in at least one direction parallel to the X axis.

2. Satellite telecommunications system (10), according to claim 1, wherein the balancing element (1284) comprises a metallic material.

3. Satellite telecommunications system (10), according to any one of the preceding claims, wherein the balancing element (1284) has a compact form.

4. Satellite telecommunications system (10), according to claim 3, wherein the compact form of the balancing element (1284) is a rectangular prism.

5. Satellite telecommunications system (10), according to claim 3, wherein the compact form of the balancing element (1284) is a cylinder.

6. Satellite telecommunications system (10), according to any one of the preceding claims, wherein the mass of the balancing element Mf, is equal to the mass of the instrument Mr divided by two, the length of the deployment device being equal to the length of the instrument L( multiplied by two).

7. Satellite telecommunications system (10), according to any one of the preceding claims, wherein the balancing element (1284) comprises at least one other payload instrument.

8. Satellite telecommunications system (10), according to any one of the preceding claims, wherein the deployment device (1282) comprises a telescopic arm.

9. Satellite telecommunications system (10), according to claim 8, wherein the telescopic arm comprises at least two sections (1282-1, 1282-2) connected by a sliding link (1282-3) having one degree of freedom, said deployment device (1282) moving the balancing element (1284) by performing a translation in a direction parallel to the X axis.

10. Satellite telecommunications system (10), according to any one of claims 1 to 7, wherein the deployment device (1282) comprises a multi-axis and multi-arm mechanism.

11. Satellite telecommunications system (10), according to claim 10, wherein the multi-axis and multi-arm mechanism comprises at least two sections (1282-4, 1282-5) connected to each other and to the body (122) by pivot links (1282-7, 1282-8) each having one degree of rotational freedom, said deployment device (1282) moving the balancing element (1284) at least in the (X, Y) plane.

12. Satellite telecommunications system (10), according to any one of the preceding claims, wherein the satellite system (10) includes a plasma-propellant propulsion device (126).

13.

14.

15. Telecommunications satellite system (10), according to any one of the preceding claims, wherein the satellite system (10) comprises a solar power generation system (124) and a control unit (1286) of the balancing device (128), said balancing device (128) being actuated by command from the control unit (1286) using the energy supplied by the solar power generation system (124). Method for actuation of the balancing device (128) of the center of inertia of a satellite telecommunications system (10) according to any one of claims 1 to 13, the method comprising the steps of: - (720) perform the launch of the satellite system (10) and po position the satellite system (10) in geostationary orbit (20); - (740) deploy the payload instrument (142) attached to the satellite system (10) by one of the two faces of the body (1222) substantially orthogonal to the X axis; - (760) deploy the balancing device (128) fixed to satellite system (10) by the face of the body (1224) substantially orthogonal to the X axis opposite said instrument fixing face (1222); the process further comprising at least one iteration of the following steps, in response to a disturbance or a maintenance operation of the satellite system (10): - (782) apply system calibration maneuvers tellitaire (10) and calculate the displacement of the center of inertia O' relative to its equilibrium position O; - (784) apply a process to define a plan of operation of the balancing device (128) for correcting the displaced center of inertia O'; - (786) operate the balancing device controls (128) correction of the center of inertia. Method for actuation of the balancing device (128) of the center of inertia of a satellite telecommunications system (10), according to the claim 14, wherein the maneuver plan is determined by a ground station and transmitted to the satellite system (10).

16. Method of actuation of the balancing device (128) of the center of inertia of a satellite telecommunications system (10), according to claim 14, wherein the maneuver plan is implemented by the satellite system (10) comprising an autonomous processing and control unit.

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