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  • 专利说明
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    • 专利摘要:
    The present invention relates to a method (50) for controlling the orbit of a satellite (10) in Earth orbit, the orbit of the satellite (10) being controlled by controlling a means of maneuver of the propulsion means (30, 31). ), comprising at least one propellant, and displacement means (20, 21) of said propulsion means. According to the invention, said maneuvering plane comprises at least two orbit control maneuvers, thrust forces of the propulsion means (30, 31) during said two orbit control maneuvers being of respective non-orbiting thrust directions. parallels in inertial reference, each of said thrust forces being determined so as to simultaneously control the inclination and longitude of the satellite orbit and to form a suitable moment to desaturate a device for storing the kinetic moment of said satellite in a plane orthogonal to the thrust direction of said thrust force.
    公开号:FR3022530A1
    申请号:FR1455630
    申请日:2014-06-19
    公开日:2015-12-25
    发明作者:Valerio Moro；Jean Fischer
    申请人:Astrium SAS；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD The present invention pertains to the field of orbit control and satellite attitude. The invention finds a particularly advantageous application, although in no way limiting, in the case of telecommunications satellites in geostationary orbit ("Geostationary Orbit" or GEO) equipped with electric propulsion means. STATE OF THE ART In known manner, a satellite in Earth orbit is subjected to numerous disturbances. These disturbances tend, on the one hand, to move the satellite relative to a set position on its orbit and, on the other hand, to change the attitude of said satellite with respect to a setpoint attitude. In order to keep the satellite substantially in the set position and in the set attitude, it is necessary to carry out an orbit control and attitude control of said satellite.
    [0002] Orbit control consists of limiting the variations of the orbital parameters generally expressed in terms of inclination, longitude and eccentricity of the satellite orbit. In the case of a GEO orbiting satellite, such as a communications satellite, orbit control is the control of the position of the satellite relative to the Earth, and is also known as satellite station keeping. ("Station keeping" or "S / K" in the Anglo-Saxon literature). Orbit control of a GEO orbiting satellite is usually accomplished by means of several orbit control maneuvers in which satellite thrusters are activated. The orbit control of the satellite is performed by adjusting the thrust forces formed by said thrusters during the various orbit control maneuvers, but also by adjusting the activation times of said thrusters. Conventionally, several orbit control maneuvers are carried out: - North / South maneuvers (N / S) make it possible to control the inclination of the satellite orbit, - East-West maneuvers (E / O) allow to control the longitude of the satellite's orbit. Eccentricity is generally controlled during maneuvers 3022530 2 E / O in the case of chemical thrusters, or during N / S maneuvers in the case of electric thrusters. It is possible to define a satellite reference center centered on a center of mass of said satellite and comprising three X, Y and Z axes: the X axis is parallel to a satellite speed vector, the Z axis is directed towards the Earth and the Y axis is orthogonal to the X and Z axes. In the satellite coordinate system, the N / S maneuvers require to have thrust forces along the Y axis, while the E / O maneuvers require to have thrust forces according to the X axis of the satellite reference. In the general case, the N / S maneuvers and the E / O maneuvers 10 use separate thrusters, which may be of different technologies (for example electrical for the N / S maneuvers and chemical for the E / O maneuvers). The thrusters used for the N / S maneuvers can be mounted on moving means. Such displacement means are used to maintain, in the YZ plane, the thrust directions of the thrusters aligned with the center of mass of the satellite (which may vary over time as a function of the amount of propellant in the engines). tanks, the position / orientation of the payload equipment, etc.), to avoid forming moments likely to change the attitude of the satellite.
    [0003] The dates of the orbit control maneuvers (i.e., the activation dates of the thrusters), the durations of said orbit control maneuvers (i.e. the activation times of the orbits). thrusters), as well as the thrust forces of said orbit control maneuvers constitute a maneuvering plane of the orbit control. This maneuvering plane is determined so as to minimize the consumption of the thrusters while maintaining the orbital parameters in predefined ranges. Attitude control involves controlling the orientation of the satellite, particularly with respect to the Earth. When the satellite is stationed in orbit, the perturbations apply moments ("torques" in the Anglo-Saxon literature) which tend to make said satellite turn around its center of mass and thus to modify the attitude of said satellite relative to to the attitude of instructions. It should be noted that the orbit control maneuvers can also apply disturbing moments when the thrust forces 3022530 3 are not perfectly aligned with the center of mass of the satellite. In order to maintain the satellite in the setpoint attitude, it is generally equipped with a kinetic moment storage device ("angular momentum" in the English literature). The kinetic moment storage device comprises, for example, at least three reaction wheels with linearly independent axes of rotation. By controlling the speed of rotation of said reaction wheels, it is possible to create moments that oppose the disturbing moments. Due to the cumulative effect of the disturbing moments, the rotational speeds of said reaction wheels, and thus the kinetic momentum stored, tend to increase gradually. It is therefore necessary to desaturate regularly the kinetic moment storage device in order to limit the speed excursion of said reaction wheels. By "desaturate" it is meant to apply external moments on the satellite which, when taken up by the kinetic moment storage device, make it possible to reduce the amount of kinetic momentum stored. Such desaturation of the kinetic moment storage device is known by the name of "angular momentum unloading" in the Anglo-Saxon literature. Desaturation of the kinetic moment storage device generally employs dedicated thrusters, which are activated during dedicated attitude control maneuvers. It will therefore be understood that the orbit control and the attitude control, in particular the desaturation of a kinetic moment storage device of a satellite, implement many different thrusters and / or many different maneuvers. Because it has many different thrusters, the complexity and cost of manufacturing the satellite are increased. Because many different maneuvers must be performed, the consumption of the thrusters is increased, which can decrease the life of the satellite, especially in the case of chemical thrusters. In addition, the increase in the number of ON / OFF sequences of the thrusters also has a negative effect on their lifetime. In addition, the operational load of the ground segment is directly related to the number of maneuvers. It is therefore desirable to limit their number.
    [0004] SUMMARY OF THE INVENTION The present invention aims 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 that makes it possible to limit both the number of 5 thrusters and the number of maneuvers required to control the satellite's orbit and desaturate a device for storing the kinetic moment of the satellite. For this purpose, and according to a first aspect, the invention relates to a method for controlling the orbit of a satellite in Earth orbit, in which the orbit of the satellite is controlled by controlling, according to a maneuvering plane, means of propulsion. , comprising at least one thruster, and means for moving said propulsion means in a satellite reference centered on a center of mass of the satellite and comprising three axes X, Y and Z, the X axis being parallel to a satellite speed vector , the Z axis being directed towards the Earth, and the Y axis being orthogonal to the X and Z axes. The displacement means 15 are furthermore suitable for: modifying angles between a thrust direction of each thruster and the axes respectively X, Y of the satellite reference, - move each thruster, constant thrust direction in the satellite reference, so as to form any axis moment 20 in a plane orthogonal to said thrust direction, the mano plane comprises at least two orbit control maneuvers, thrust forces of the propulsion means during said two orbit control maneuvers are respective non-parallel, inertial reference thrust directions, and each of said thrust forces is determined so as simultaneously to control the inclination and the longitude of the orbit of the satellite and to form a moment suitable for desaturating a device for storing the kinetic moment of said satellite in a plane orthogonal to the direction of thrust of said force of thrust. In general, throughout the present application, a thrust force is defined by a thrust vector and a point of application of said thrust force relative to the center of mass of the satellite. The thrust vector is itself defined by a thrust standard and a unit standard thrust direction, which corresponds to the thrust vector 3022530 normalized by said thrust standard. As indicated above, the displacement means make it possible to modify angles between the direction of thrust of each thruster and the axes X, Y respectively of the satellite reference. Therefore, it is possible with such moving means to form a thrust direction thrust force adapted to simultaneously control the inclination of the satellite orbit (component along the Y axis) and the longitude of the orbit of the satellite (component along the X axis). The displacement means further enable each thruster, with constant thrust direction, to be moved in the satellite fixture (i.e. only the point of application of the thrust force is displaced), so as to form a moment of any axis in a plane orthogonal to said direction of thrust. Therefore, it is possible with such moving means to form a thrust force adapted not only to control the inclination and longitude of the satellite orbit, but also to form a desaturation moment of the satellite storage device. kinetic moment included in the plane orthogonal to said direction of thrust. Indeed, for a given thrust direction, the possible axes of the moment formed by moving the point of application of the thrust force are all in the plane orthogonal to said thrust direction. Therefore, for a given thrust direction, the desaturation of the kinetic moment storage device is only possible in a two dimensional vector space. Since the maneuvering plane comprises at least two orbit control maneuvers for which the thrust forces are of non-parallel respective thrust directions in an inertial frame, the planes orthogonal to these thrust directions are also not parallel to each other. so that, on average on these two orbit control maneuvers, the desaturation of the kinetic moment storage device is possible in a vector space of dimension three.
    [0005] In particular embodiments, the satellite orbit control method may further include one or more of the following features, taken alone or in any technically feasible combination.
    [0006] In particular embodiments, the maneuvering plane is determined under the constraint: ## EQU1 ## where: 5 - r is a strictly positive scalar value, ENI is the ratio between an X-axis component and a Y-axis component of the thrust force of a first maneuver control maneuver of the maneuvering plane, - EN2 is the ratio of a component following the X axis 10 and a component along the Y axis of the thrust force of a second orbit control maneuver maneuver plane, - RN corresponds to the ratio between a component along the Z axis and the component following the Y axis of the thrust force of the first or second control maneuver maneuver plane, -AT is equal to 27. (T2 - Ti - Torb / 2) / Torb, expression in which Ti and T2 are dates of the first and second orbit control maneuvers, and Torb is the period o satellite. Such arrangements are advantageous in that they make it possible to determine a maneuvering plane ensuring a minimum value r of the minimum three-axis desaturation capacity of the kinetic moment storage device. For example, it is possible to dynamically adapt said minimum value r as a function of the kinetic moment stored in the kinetic moment storage device. Indeed, if the thrust forces of the at least two maneuver control maneuvers of the maneuvering plane are of almost parallel thrust directions in an inertial frame, then the moment-forming capability in a direction of average thrust of said forces thrust will be weak. If it is found that it is not possible to effectively desatter the kinetic moment storage device, then this will indicate that it is necessary to increase the moment-forming capability in said average thrust direction, which can be simply taken into account in the determination of the maneuvering plane by increasing the minimum value r of the desaturation capacity 3022530 7 three axes. In particular embodiments, the displacement means comprising an articulated arm carrying a propulsion means propulsion, said articulated arm having at least three joints 5 each having at least one degree of freedom in rotation about an axis of rotation, the respective axes of rotation of adjacent joints being not parallel for at least two pairs of adjacent joints, the thrust force of said thruster, in particular the thrust direction and the point of application of said thrust force , is controlled by controlling the articulations of the articulated arm. In particular embodiments, the displacement means comprising a satellite attitude control device and an articulated arm carrying a thruster of the propulsion means, said articulated arm comprising at least two joints each comprising at least one degree. of freedom in rotation, the thrust force of the thruster is controlled by controlling the joints of the articulated arm and the attitude of the satellite. In particular embodiments, the articulated arm comprising at least one additional articulation, is controlled, during at least one orbit control maneuver maneuvering plane, the means 20 for propulsion and the means for moving the so as to form a steering thrust force adapted to further control the eccentricity of the orbit of the satellite. In particular modes of implementation, the eccentricity of the satellite orbit is controlled by controlling additional propulsion means of said satellite, fixed orientation relative to said satellite. In particular embodiments, the dates and / or durations of the two maneuvering orbit control maneuvers are determined so as to further control the eccentricity of the satellite orbit in the course of the maneuver.
    [0007] In particular modes of implementation, an intermediate maneuvering plane, adapted to control only the orbit of the satellite, is determined by a terrestrial station and transmitted to the satellite, and the maneuvering plane to be implemented is determined by the satellite according to said intermediate plane 3022530 8 to also perform the desaturation. In particular modes of implementation, the maneuvering plane comprises at most two orbit control maneuvers per orbital period of the satellite.
    [0008] According to a second aspect, the present invention relates to a computer program product characterized in that it comprises a set of program code instructions which, when executed by a processor, implement a control method. orbit of a satellite according to any one of the embodiments of the invention.
    [0009] According to a third aspect, the present invention relates to a satellite intended to be deployed in terrestrial orbit, comprising propulsion means, comprising at least one propellant, and means for moving said propulsion means in a satellite reference center centered on a center of mass of said satellite and comprising three axes X, Y and Z, the X axis being parallel to a satellite velocity vector, the Z axis being directed towards the Earth, and the Y axis being orthogonal to the X and Z. In addition, the displacement means are suitable for: modifying angles between a thrust direction of each thruster and the axes X, Y respectively of the satellite reference 20 - moving each thruster, with a constant thrust direction in the satellite reference , so as to form a moment of any axis in a plane orthogonal to said direction of thrust. The satellite further comprises means adapted to control the propulsion means and the displacement means according to a maneuvering plane comprising at least two orbital control maneuvers, thrust forces of the propulsion means during said two maneuvers. orbit control being of non-parallel respective thrust directions in an inertial frame, each of said thrust forces being determined so as to simultaneously control the inclination and longitude of the satellite orbit and to form a suitable moment for desaturating a device for storing the kinetic moment of said satellite in a plane orthogonal to the thrust direction of said thrust force. In particular embodiments, the satellite may further include one or more of the following features, taken alone or in any technically feasible combination. In particular embodiments, the displacement means are arranged non-symmetrically with respect to the XZ plane 5 formed by the X and Z axes of the satellite reference. Such arrangements make it possible to facilitate the desaturation of the kinetic moment storage device. In particular embodiments, the displacement means comprise an articulated arm carrying a propulsion means of propulsion, said articulated arm having at least three joints 10 each having at least one degree of freedom in rotation about an axis of rotation, the respective axes of rotation of adjacent joints not being parallel to each other for at least two pairs of adjacent joints. In particular embodiments, the articulated arm comprises an additional articulation comprising at least one degree of freedom in rotation about an axis of rotation. In particular embodiments, the satellite comprises additional propulsion means, fixed orientation relative to said satellite. In particular embodiments, the propulsion means carried by the displacement means are electric propulsion means. According to a fourth aspect, the present invention relates to an orbit control system of a satellite according to any one of the embodiments of the invention, comprising means adapted to determine the maneuvering plane under the constraint: IEN1 + EN2 + RN.sin (AT) I> r where: - r is a strictly positive scalar value, - ENI is the ratio of a component along the X axis to a component along the Y axis of the force of pushing of a first maneuver control maneuver of the maneuvering plane, - EN2 corresponds to the ratio between a component along the X axis and a component along the Y axis of the pushing force of a second control maneuver of the maneuvering plane, RN is the ratio between a Z-axis component and the Y-axis component of the thrust force of the first or second orbital control maneuver. maneuvering plane, 5 - AT is equal to 27. ( T2 - T1 - Torb / 2) / Torb, where T1 and T2 are dates of the first and second orbit control maneuvers, and Torb is the orbital period of the satellite. In particular embodiments, the satellite orbit control system may further include one or more of the following features, taken alone or in any technically possible combination. In particular embodiments, the means adapted to determine the maneuvering plane are distributed between the satellite and a terrestrial station.
    [0010] In particular embodiments, an intermediate maneuvering plane, adapted to control only the orbit of the satellite, is determined by the earth station and transmitted to the satellite, and the maneuvering plane to be implemented is determined by the satellite. function of said intermediate maneuvering plane.
    [0011] PRESENTATION OF THE FIGURES The invention will be better understood on reading the following description, given by way of non-limiting example, and with reference to the figures which represent: FIG. 1: a schematic representation of a system of FIG. orbit control of a satellite orbiting the Earth, - Figure 2: a schematic representation of a particular embodiment of a satellite according to the invention, - Figure 3: a schematic representation of a preferred mode embodiment of a satellite according to the invention, FIG. 4: a schematic representation of an alternative embodiment of the satellite of FIG. 3; FIG. 5: a diagram representing the main steps of a control method of FIG. orbit of a satellite according to the invention.
    [0012] In these figures, identical references from one figure to another designate identical or similar elements. For the sake of clarity, the elements shown are not to scale unless otherwise stated. DETAILED DESCRIPTION OF EMBODIMENTS FIG. 1 schematically represents an orbit control system of a satellite 10. In the remainder of the description, reference is made in a nonlimiting manner to the case of a satellite 10 in GEO orbit. . However, there is nothing to preclude, according to other examples, consideration of other types of spacecraft (space shuttle, orbital station, etc.), and / or other terrestrial orbits, for example geosynchronous orbits, medium orbits ("Medium Earth Orbit" or MEO), low orbits ("Low Earth Orbit" or LEO), etc. For the purposes of the description, the satellite 10 is associated with a satellite reference centered on a center of mass O of the satellite 10 and having three X, Y, Z axes. More particularly, the X axis is parallel to a vector of the X axis. satellite 10 in inertial reference, the Z axis is directed towards the center of the Earth T, and the Y axis is orthogonal to the X and Z axes. Each of the X, Y and Z axes of the satellite coordinate system is associated with unit vectors. respectively ux, uy and uz. The unit vector ux corresponds to the velocity vector normalized by the norm of said velocity vector, the unit vector uz is oriented from the center of mass O of the satellite 10 to the center of the Earth T, and the unit vector uy is oriented so that the set (ux, uy, uz) constitutes a direct orthonormal basis of the satellite reference. As illustrated in FIG. 1, the satellite 10 comprises, for example, a body 11, and two solar generators 12 on either side of the body 11. The two solar generators 12 are for example mounted to rotate relative to each other. body 11 of the satellite 10, around the same axis of rotation. In the remainder of the description, one places oneself in a nonlimiting manner in the case where the body 11 of the satellite 10 is substantially in the form of a rectangular parallelepiped. The body 11 thus has six two by two parallel faces, and the two solar generators 12 are respectively arranged on two opposite faces of said body 11, the axis of rotation of said two solar generators 12 being substantially orthogonal to said two opposite faces 3022530 12 of the body 11 of the satellite 10. In the remainder of the description, one places oneself in a nonlimiting manner in the case where the attitude of the satellite 10 is controlled, for the purposes of the mission of said satellite 10, so as to be placed in an attitude setpoint, 5 called "mission attitude", wherein: - a face of the body 11 of the satellite 10, designated by "face + Z", carrying for example an instrument of a payload of said satellite 10, is directed towards the Earth and is substantially orthogonal to the Z axis; the opposite face to the + Z face, then arranged on the opposite side of the Earth, is designated "-Z" face; the two opposite faces of the body 11 of the satellite 10 on which are arranged the two solar generators 12, designated respectively by "face + Y" (with respect to the center of mass O: the side pointed by the unit vector uy) and "face -Y "are substantially orthogonal to the Y axis; the last two opposite faces of the body 11 of the satellite 10, designated respectively by "face + X" (with respect to the center of mass O: on the side pointed by the unit vector ux) and "face X", are substantially orthogonal to the X axis.
    [0013] The satellite 10 also comprises a set of actuators adapted to control the orbit and the attitude of said satellite 10, as well as a control device (not shown in the figures) of said actuators. For the purposes of attitude control, the satellite 10 comprises in particular a kinetic moment storage device (not shown in the figures) adapted to store a kinetic moment of any axis, that is to say having a kinetic moment storage capacity along three linearly independent axes. The kinetic moment storage device comprises a set of inertial actuators such as reaction wheels and / or gyroscopic actuators. For example, the kinetic moment storage device comprises at least three respective linearly independent rotational axis reaction wheels. As indicated above, the orbit control consists of controlling at least one orbital parameter among the inclination, the longitude and the eccentricity 3022530 13 of the orbit of the satellite 10. In the case of a satellite 10 in orbit GEO, it is known that the requirements in terms of orbit control, for example expressed in terms of required speed variation per year (m / s / yr), are mainly imposed by the control of the inclination of the orbit satellite 10 (N / S control).
    [0014] The order of magnitude of the required speed variation per year for the N / S control, along the Y axis, is thus 50 m / s / year, while it is 2-4 m / s / year for the control of the longitude of the orbit (control E / O), along the axis X. For the purposes of the orbit control, the satellite 10 comprises in particular propulsion means, comprising at least one propellant, and means for moving said propulsion means in the satellite reference. More particularly, the displacement means are suitable for: modifying angles between a thrust direction of each thruster and the axes X, Y respectively of the satellite marker; moving each thruster, with a constant thrust direction, in the satellite reference, so as to form a moment of any axis in a plane orthogonal to said direction of thrust (including a zero moment by aligning the thrust direction with the center of mass O of the satellite 10). The orbit control of the satellite 10 is carried out at the level of the control device 20 by controlling the propulsion means and the displacement means according to a maneuvering plane comprising orbit control maneuvers during which the propulsion means are activated. With such displacement means, it is understood that it is possible, during the same orbit control maneuver and with the same thruster, to control the thrust direction of said thruster so as to simultaneously control the inclination (adjusting the Y-direction component of the thrust direction) and longitude (adjusting the X-direction direction of the thrust direction component) of the orbit. It is also possible, again during the same orbit control maneuver and with the same thruster, to form if necessary a moment of desaturation of the kinetic moment storage device along any axis included in the plane orthogonal to said thrust direction, by moving the point of application of the thrust force relative to the center of mass of the satellite 10.
    [0015] In order to be able to desaturate the kinetic moment storage device irrespective of the direction of the kinetic moment stored, the maneuvering plane advantageously comprises at least two orbit control maneuvers whose respective thrust forces of the propulsion means are non-parallel inertial reference. Thus, during said two orbit control maneuvers, the planes in which it is possible to form a moment of desaturation are not parallel so that, on all of said two orbital control maneuvers, space vector in which it is possible to form a moment of desaturation is of dimension three.
    [0016] The thrust forces of the maneuvering plane are therefore determined, according to predetermined needs for controlling the inclination and the longitude of the orbit of the satellite 10, and according to a predetermined need for desaturation of the storage device. of kinetic moment of said satellite 10, so as to simultaneously control the inclination and longitude of the orbit while desaturating the kinetic moment storage device of said satellite 10. In the prior art, the E / O maneuvers were performed at a lower frequency than the N / S plasma maneuvers. According to the invention, the E / O control is carried out simultaneously with the N / S control. Therefore, the frequency of the E / O control is greater than the frequency of the E / O maneuvers of the prior art, so that it is possible for example to reduce the excursion in longitude relative to that of the I / O control. prior art. In addition, the number of orbit control maneuvers is reduced compared to the prior art in that the inclination and longitude of the orbit of the satellite 10 are simultaneously controlled. The number of activations of the propulsion means is therefore also reduced and the consumption, in particular in the case of chemical propulsion means, can be reduced. Finally, the total number of orbit control and desaturation maneuvers of the kinetic moment storage device is also reduced compared to the prior art in that it is no longer necessary to have dedicated maneuvers of desaturation. In addition, the same propulsion means are used both to control the orbit of the satellite 10 and to desaturate the kinetic moment storage device, so that it is possible to reduce the number of on-board thrusters. However, nothing precludes the provision of additional propulsion means dedicated to the desaturation of the kinetic moment storage device. If necessary, it will be possible to reduce the capacity of the said additional propulsion means and / or the kinetic moment storage device, to the extent that the propulsion means used to control the orbit of the satellite 10 also contribute to regularly desaturing said kinetic moment storage device. In preferred embodiments, it is also possible, again during the same orbit control maneuver, to further control the eccentricity of the orbit, for example by adjusting the dates and / or durations of said at least two maneuver orbit control maneuvering plan according to a predetermined need for control of the eccentricity of the orbit. In such embodiments, all the orbital parameters are therefore controlled simultaneously, and this further desaturates the kinetic moment storage device. Preferably, in the case of a satellite in GEO orbit (or more generally: in geosynchronous orbit), the maneuvering plane comprises altogether two orbital control maneuvers per orbital period (approximately 24 hours). For the purpose of orbit tilt control, the nominal time spacing of the orbit control maneuvers is approximately 12 hours. However, the optimal thrust directions for correcting the prevailing disturbances (inclination) being oriented according to the orbital normal, the thrust forces determined solely for the disturbance control would have substantially parallel thrust directions in inertial reference. Consequently, in preferred embodiments, a misalignment between said thrust directions and / or a time offset between the orbit control maneuvers with respect to the time spacing of the prior art is imposed. FIG. 2 shows a particular embodiment of a satellite 10 comprising displacement means adapted to simultaneously control the inclination and the longitude of the orbit of the satellite 10, while at least partially desaturating the moment storage device. kinetic. For the sake of clarity in FIG. 2, the solar generators 12 of the satellite 3022530 are not shown. In the example illustrated in FIG. 2, the displacement means comprise two articulated arms 20, 21, each articulated arm 20, 21 carrying a thruster 30, 31. In the remainder of the description, non-limiting in the case where the thrusters 30, 31 are electric thrusters (electrothermal, electrostatic, plasma, etc.). Nothing, however, excludes, according to other examples, that one or both thrusters 30, 31 are chemical propellants (cold gas, liquid propellants, etc.). In the nonlimiting example illustrated in FIG. 2, the articulated arms 20, 21 are arranged respectively on the + Y face and the Y-face of the body 11 of the satellite 10. The articulated arms 20, 21 are, for example, respectively for the South control and the North control of the inclination of the orbit of the satellite 10, by alternately activating either the thruster 30 or the thruster 31.
    [0017] Preferably, the articulated arm 20 is attached to said + Y face at a fixed point which substantially corresponds to the orthogonal projection of a theoretical center of mass of the satellite 10 on said + Y face. The theoretical center of mass (considered merged with the actual center mass O in FIG. 2) corresponds, for example, to an estimate before launch of the center of mass 20 of the satellite 10 stationed in the GEO orbit. In other words, the attachment point of the articulated arm is such that the moment applied to the satellite 10 is substantially zero when the articulated arm 20 is deployed substantially orthogonal to the + Y face by imposing a thrust direction of the thruster 30 substantially orthogonal to said + Y face. If the satellite 10 is furthermore in the mission attitude, then said thrust direction is orthogonal to the orbit plane coincident with the XZ plane, and thus only controls the inclination of the orbit. Similarly, the articulated arm 21 is attached to said Y-face at a fixed point which substantially corresponds to the orthogonal projection of the theoretical mass center of the satellite 10 on said Y-face. Thus, the moment applied to the satellite 10 is substantially zero when the articulated arm 21 is deployed substantially orthogonal to the face -Y by imposing a thrust direction of the thruster 31 substantially orthogonal to said face -Y. If the satellite 10 is further in the mission attitude, then said thrust direction is orthogonal to the orbital plane coincident with the XZ plane, and therefore only controls the inclination of the orbit. In the remainder of the description, the term "5 N / S control position" designates the position of the articulated arm 20 (respectively articulated arm 21) in which said articulated arm is deployed substantially orthogonally to the + Y face (respectively face -Y ) by imposing a thrust direction of the thruster 30 (respectively thruster 31) substantially orthogonal to said face + Y (respectively face -Y), so that the thrust force of said thruster 30 (respectively propellant 31) is substantially aligned with the center of theoretical mass of the satellite 10. In the example illustrated in FIG. 2, each articulated arm 20, 21 comprises three articulations 22, 23, 24, each articulation comprising at least one degree of freedom in rotation about an axis of rotation. The hinges 22 and 23 are interconnected by a connection 25, while the hinges 23 and 24 are interconnected by a link 26. In addition, for each hinged arm 20, 21, the respective axes of rotation of the hinges 22, 23, 24 are not parallel for each of the two pairs of adjacent joints.
    [0018] Thus, each articulated arm 20, 21 offers three degrees of freedom to modify, with respect to the N / S control position, the thrust direction and the point of application of the thrust force of the thruster 30, 31. For example, a first degree of freedom can be used to control the X-axis directional direction component (I / O control), and the other two degrees of freedom can be used to control the position of the point of rotation. application of the thrust force relative to the center of mass O of the satellite 10 (desaturation of the kinetic moment storage device). Preferably, and as illustrated in the example of FIG. 2, the axis of rotation of the articulation 22 of each articulated arm 20, 21 is, when the satellite 10 is in the mission attitude, substantially parallel to the axis Z. The axis of rotation of the joint 23 of each articulated arm 20, 21 is substantially orthogonal to both the connection 25 and the axis of rotation of the joint 22. The axis of rotation of the articulation 24 of each articulated arm 3022530 18 20, 21 is it substantially orthogonal to both the connection 26 and the axis of rotation of the articulation 23. To control the direction of thrust and the point application of the thrust force, the control device controls the angles of rotation 5 of the joints 22, 23, 24, designated respectively by 01, 02 and 03. The joints 22, 23, 24 are for example such that, when the satellite 10 is stationed on the GEO orbit, each of the angles of rotation and 02 and 03 of each arm articulated 20, 21 can take any value in a range of values [-30 °, 30 °] around the N / S control position of said articulated arm.
    [0019] In particular embodiments, and as shown in FIG. 2, the satellite 10 comprises additional propulsion means, of fixed orientation relative to said satellite 10. For example, the satellite 10 comprises a propellant 40 (chemical or electrical) fixed to the face -Z of the body 11 of the satellite 10, of fixed orientation such that the thrust direction of said thruster 40 is substantially orthogonal to said -Z face. Preferably, the point of attachment of the thruster 40 to the -Z face substantially corresponds to the orthogonal projection of the theoretical center of mass of the satellite 10 on said -Z face. Thus, the moment applied to the satellite 10 by said thruster 40 is substantially zero as long as the actual center of mass O of the satellite 10 is close to the theoretical center of mass. It should be noted that the satellite 10 may comprise, according to other examples, several thrusters 40 fixed orientation relative to the satellite 10. The thruster 40 is implemented to control the eccentricity of the orbit. It can be activated simultaneously with the thrusters 30, 31 carried by the articulated arms 20, 21, and / or during dedicated eccentricity control maneuvers, distinct from the control operations N / S and E / O of the orbit. . In addition to or alternatively with the thruster 40 having a fixed orientation relative to the satellite 10, at least one of the articulated arms 20, 21 may comprise, in particular embodiments, an additional articulation 30 (not shown in the figures). having at least one degree of freedom in rotation about an axis of rotation. This additional articulation is for example connected to the articulation 24 by an additional connection, and the axis of rotation of said additional articulation is preferably orthogonal to the axis of rotation of the articulation 24 and to said additional connection. Each articulated arm 20, 21 comprising such additional articulation then comprises an additional degree of freedom, which can for example be used by the control device to control all the orbital parameters, including the eccentricity, simultaneously with the control maneuvers. / S and E / O of the satellite orbit 10. FIG. 3 represents a preferred embodiment of a satellite 10 which comprises the same displacement means (articulated arms 20 and 21) and the same propulsion means (thrusters 30 and 31) carried said displacement means 10 as the satellite 10 illustrated in FIG. 2. As illustrated by FIG. 3, the articulated arm 20 is advantageously fixed to the + Y face at a fixed point which is shifted, following the Z axis, relative to the orthogonal projection of the theoretical center of mass of the satellite 10 on said + Y face. Similarly, the articulated arm 21 is advantageously fixed to the face Y at a fixed point which is offset along the axis Z relative to the orthogonal projection of the theoretical center of mass of the satellite 10 on said face Y . In the example illustrated in FIG. 3, the N / S control position of the articulated arm 20 (respectively articulated arm 21) corresponds to the position in which the link 25 is oriented substantially orthogonally to the + Y face 20 (respectively face - Y) and the link 26 is oriented substantially towards the theoretical center of mass of the satellite 10, by imposing a thrust force of the thruster 30 (respectively thruster 31) substantially aligned with said theoretical center of mass of the satellite 10. Such a configuration of articulated arms 20, 21, compared with that illustrated in Figure 2, facilitates the control of the eccentricity of the orbit of the satellite 10, insofar as the thrust force of the thruster 30 (propellant 31 respectively) , in the N / S control position, comprises a non-zero component along the Z axis without forming a moment. It is therefore possible, with the only articulated arms 20, 21 and the only thrusters 30, 31 to control all the orbital parameters, including the eccentricity of the orbit of the satellite 10, for example by adjusting the times and / or the dates of the control maneuver maneuvers of the maneuvering plane. The satellite 10 can therefore, as is the case in the nonlimiting example illustrated in FIG. 3, be devoid of a fixed orientation propulsion unit 40. However, as shown in FIG. 4, there is nothing to preclude the satellite 10 of FIG. 3 from being equipped with a thruster 40 of fixed orientation, for purposes such as redundancy to overcome a failure of one of the thrusters 30, 31.
    [0020] In the examples illustrated in FIGS. 2, 3 and 4, the displacement means, that is to say the articulated arms 20, 21, are arranged symmetrically with respect to the XZ plane formed by the X and Z axes. satellite reference. In particular, in these examples, the attachment points of the articulated arms 20, 21 are arranged substantially on the same axis parallel to the axis Y, so that the coordinates along the X and Z axes of said attachment points of said articulated arms 20, 21, in a satellite reference centered on the theoretical center of mass, are both identical. In preferred alternative embodiments, not shown in figures, the articulated arms 20, 21 are arranged non-symmetrically with respect to the XZ plane. In particular, the coordinates along the X and Z axes of said attachment points of said articulated arms 20, 21, in the satellite coordinate system centered on the theoretical center of mass, are preferably not both identical. Such arrangements make it possible to facilitate the desaturation of the kinetic moment storage device of the satellite 10.
    [0021] As discussed above, the displacement means comprise, in the examples illustrated in Figures 2, 3 and 4, two articulated arms 20, 21 each having at least three joints 22, 23, 24. However, other Embodiments of the displacement means are possible, without the invention being modified in principle. In particular, nothing makes it possible, according to other embodiments, to have articulated arms comprising two articulations each comprising at least one degree of freedom in rotation about an axis of rotation, the axes of rotation of said two joints are not parallel. In such a case, each articulated arm has two degrees of freedom, and it is possible to obtain an additional degree of freedom by changing the attitude of the satellite 10 in the satellite fixture during the orbit control maneuvers. Where appropriate, the displacement means further comprise a satellite attitude control device 10, which may be the kinetic moment storage device.
    [0022] In addition, a satellite 10 according to the invention has been described with reference to FIGS. 2, 3 and 4, considering that the displacement means comprise two articulated arms 20, 21. Nothing excludes, according to other examples, to have a number of articulated arms different from two. In particular, the displacement means may comprise only one articulated arm. In the case of a single articulated arm, said articulated arm is for example fixed to the face -Z of the body of the satellite 10, or to one of the two faces + Y and -Y, as illustrated by FIGS. , 3 and 4. In the case of a single articulated arm attached to the face -Y (or face + Y) of the body 11 of the satellite 10, it is possible, for example, for the South control (respectively the control North) of the orbit of the satellite 10, to rotate the satellite 10, with respect to the mission attitude, 180 ° around the Z axis of the satellite reference. As previously indicated, the control device controls the orbit of the satellite 10 as a function of a maneuvering plane comprising at least two orbit control maneuvers with non-parallel reference thrust forces inertial. In addition, each of said thrust forces is determined so as to simultaneously control the inclination and longitude (and possibly the eccentricity) of the satellite orbit 10 as well as to form a suitable moment to desaturate the storage device. the kinetic moment of said satellite 10. The control device comprises for example at least one processor and at least one electronic memory in which is stored a computer program product, in the form of a set of program code instructions to execute to control the means of displacement and the propulsion means of the satellite 10 according to such a maneuvering plane. In one variant, the control device comprises one or more programmable logic circuits, of the FPGA, PLD, etc. type, and / or specialized integrated circuits (ASIC) adapted to implement all or part of said control steps of the displacement means. and propulsion means according to such a maneuvering plane. In other words, the control device comprises a set of means configured in software (specific computer program product) and / or hardware (FPGA, PLD, ASIC, etc.) for controlling the movement means and the propulsion means of the satellite 10 according to such a maneuvering plane. The main parameters to be adjusted of the maneuvering plane are: the start dates of the various orbital control maneuvers of the maneuvering plane, that is to say the activation dates of the propulsion means, the durations various control maneuvering maneuvers of the maneuvering plane, that is to say the activation times of the propulsion means, 10 - the thrust directions and the points of application relative to the center of mass O of the satellite 10 respective pushing forces of the various orbit control maneuvers. In the remainder of the description, reference is made in a non-limiting manner in the case where the thrust direction and the point of application of each thrust force of the maneuvering plane are fixed relative to the satellite 10 for the duration of the the corresponding orbit control maneuver. In other words, in the case of the displacement means illustrated in FIGS. 2, 3 and 4, the values of the angles O1, O2 and O3 of the articulations 22, 23, 24 of the articulated arms 20, 21 are not modified at during the same orbit control maneuver. However, according to other examples, nothing makes it possible to vary said values of the angles 01, 02 and 03 in order to increase the number of degrees of freedom of the orbit control system. It is also possible to adjust other parameters such as the respective durations of the different maneuvering orbit control maneuvers and / or the pushing norms of the respective pushing forces of said different orbit control maneuvers ( in the case of propulsion means whose thrust standard can be controlled). F1 denotes the thrust force of the propulsion means during the first of the two orbit control maneuvers, which starts at a date T1, and by F2 the thrust force of the propulsion means during the second the two orbital control maneuvers, which starts on a T2 date. The pushing forces F1 and F2 are expressed in the satellite coordinate system according to the following expressions: ## EQU1 ## F1 (F1) + F1z.uz (T1) F2 = F2x-ux (T2) ) + F2y.uy (T2) + F2z-uz (T2) expression in which: - (F1x, F1y, F1z) are the components of the thrust force F1 5 in the satellite reference at the date T1, whose unit vectors are (ux (T1), uy (T1), uz (T1)), - (F2x, F2y, F2z) are the components of the pushing force F2 in the satellite coordinate system at the date T2, whose unit vectors are ( ux (T2), uy (T2), uz (T2)).
    [0023] 10 If one places oneself in the case of a satellite 10 as illustrated by FIG. 3 in which the eccentricity control is carried out by means of the thrusters 30, 31 carried by the articulated arms 20, 21, then the The system of equations to be solved has nine equations, relating to the following parameters: - AVx (T1) and AVx (T2), scalar parameters which correspond to the needs in terms of required speed variation along the X axis (control E / O), during the first maneuver of orbit control and the second maneuver of the maneuvering plane, - AVy (T1) and AVy (T2), scalar parameters which correspond to the needs in terms of speed variation required following the axis 20 Y (N / S control), during the first orbit control maneuver and the second maneuver of the maneuvering plane, - AVz (T1) and AVz (T2), scalar parameters which correspond to the needs in terms of required speed variation along the Z axis (eccentric control during the first orbit control maneuver and the second maneuver of the maneuvering plane, AH, a vector of three scalar parameters corresponding to the components of the kinetic moment to be destocked from the kinetic moment storage device. the outcome of the two orbit control maneuvers, expressed in inertial reference. By adjusting the durations and dates of the two orbit control maneuvers and the angle values 81 (T1), 82 (T1), 83 (T1), 81 (T2), 82 (T2) and 83 (T2) 22, 23, 24 joints articulated arms 20, 21 during said two orbit control maneuvers, there is then a sufficient number of degrees of freedom to solve the system of equations mentioned above. If we consider the case of a satellite 10 in which the eccentricity control is carried out by means of a thruster 40 of fixed orientation relative to the satellite 10, and in the case where said satellite 10 is without one of the two thrusters 30 or 31 (by design or because of a temporary or permanent failure of said thruster or the articulated arm carrying it), then the loss of the associated degrees of freedom can be compensated by an increase in the number of orbit control maneuvers. Preferably, the maneuvering plane is in this case executed over a longer time horizon, so as not to increase the number of orbital control maneuvers per orbital period. For example, it is possible to consider a maneuvering plan spanning several orbital periods, preferably comprising at most two orbital control maneuvers per orbital period. The advantages are that the operational load on the ground is unchanged after the failure, and that there is no increase in the number of ON / OFF sequences of the thrusters (important with respect to the service life, especially after a breakdown). The operation with a greater frequency of maneuvers, to have a better accuracy on the orbit control, is nevertheless nevertheless possible. In preferred embodiments, the maneuvering plane is further determined under the following constraint: IEN1 + EN2 + RN.sin (OT) I> r (1) where: 25 - r is a scalar value strictly positive representative of a minimum required three-axis desaturation capacity of the kinetic moment storage device of the satellite 10, - EN1 is equal to the ratio F1x / F1y, - EN2 is equal to the ratio F2x / F2y, 30 - RN is equal to the ratio F1z / F1y or the ratio F2z / F2y, - AT is equal to 27. (T2 - T1 - Torb / 2) / Torb, where Torb is the orbital period (approximately 24 hours in the case of a geosynchronous orbit) .
    [0024] It should be noted that other constraints can also be considered, in addition to the constraint given by the expression (1), during the determination of the maneuvering plane. AT is therefore representative of the time offset with respect to the nominal time spacing (Torb / 2, ie 12 hours in the case of a geosynchronous orbit) between the orbit control maneuvers. If AT is non-zero modulo Tr, then the thrust forces F1 and F2, determined solely for the control of the inclination and longitude of the orbit, will be nonparallel. In addition, the ratios EN1 and EN2 make it possible to impose misalignment between said thrust forces F1 and F2 (by imposing that the sum EN1 + EN2 is not zero). Finally, the global expression (1) above makes it possible to ensure that, if one imposes both a misalignment of the thrust forces F1 and F2 in satellite reference and a temporal offset with respect to the nominal time spacing, those they do not cancel each other out in inertial reference.
    [0025] The value r is, for example, a constant predefined value over time, or a value that can be adjusted over time, depending for example on the kinetic moment AH to be destocked. In particular, it is understood that, if the thrust forces F1 and F2 are almost parallel in inertial reference, then the desaturation capacity is low in the direction of average thrust of said thrust forces F1 and F2. The higher the value r increases, and the greater the absolute value of the scalar product of the thrust directions of the thrust forces F1 and F2 in the inertial frame tends to decrease, so that the three-axis desaturation capacity of the kinetic moment storage device increases.
    [0026] The maneuvering plane is for example determined directly by the satellite control device 10. Alternatively, the maneuvering plane can be determined by a terrestrial station of the orbit control system, and transmitted to the satellite 10 to be implemented. by the control device. The maneuvering plane can also, according to other examples, be jointly determined by the satellite 10 and the earth station. FIG. 5 schematically represents the main steps of a preferred embodiment of an orbit control method 50 in which the maneuvering plane is jointly determined by a terrestrial station and by the satellite 10. For this purpose, the land station and the satellite 10 comprise respective conventional communication means which enable them to exchange data. As illustrated in FIG. 5, the orbit control method 50 firstly comprises a step 52 for determining an intermediate maneuvering plane, comprising two orbit control maneuvers whose thrust forces, non-parallel thrust directions are adapted to simultaneously control several orbital parameters (tilt, longitude, and possibly eccentricity) of the satellite orbit 10 without changing its kinetic moment. For example, during step 52 of determining the intermediate maneuvering plane, the earth station determines the following parameters: the dates T1 and T2 of the start of the orbit control maneuvers, the durations of the two maneuvers of orbit control, - intermediate values of the angles 81 (T1), 82 (T1), 83 (T1), 61 (T2), 82 (T2) and 83 (T2) making it possible to have thrust forces which control simultaneously the inclination and longitude of the orbit of the satellite 10 without forming a moment.
    [0027] These parameters are for example determined according to the scalar parameters AVx (T1), AVx (T2), AVy (T1), AVy (T2) (and possibly AVz (T1) and AVz (T2)), received from the satellite 10 or determined directly by the ground station. Preferably, the intermediate maneuvering plane is determined by the earth station under the constraint IEN1 + EN2 + RN.sin (AT) I> F. If the value F is not constant over time, this is for example received from the satellite 10, or determined from data received from the satellite 10, such as the kinetic moment AH destocking. The intermediate maneuvering plane, once determined by the earth station, is transmitted to the satellite 10. The orbit control method 50 then comprises a step 54 during which the control device determines the maneuvering plane to be implemented. implement, that is to say the operating plane whose thrust forces 3022530 27 are further adapted to form desaturation moments of the kinetic moment storage device. More particularly, this step 54 aims at determining new values of the angles O1 (T1), O2 (T1), O3 (T1), O1 (T2), O2 (T2) and O3 (T2), making it possible in addition to destock a kinetic momentum AH of the kinetic moment storage device at the end of the two orbit control maneuvers. Said new values of the angles O1 (T1), O2 (T1), O3 (T1), O1 (T2), O2 (T2) and O3 (T2) are determined as a function of the kinetic moment AH to be destocked and as a function of the intermediate values. said angles given by the intermediate maneuvering plane, for example by linearizing the system of equations around said intermediate values. The orbit control method 50 then comprises a step 56 during which the control device controls the displacement means and the propulsion means in accordance with the maneuvering plane determined during the step 54. As indicated previously, the maneuvering plan to be implemented can also be entirely determined by the ground station. Where appropriate, the step 54 of determining the maneuvering plane to implement, in particular according to the intermediate maneuvering plane, is performed by the ground station. Nothing further excludes, according to other examples, to determine then directly the maneuver plan to implement, without going through the determination of an intermediate maneuvering plan. The terrestrial station comprises, for example, at least one processor and at least one electronic memory in which a computer program product is stored in the form of a set of program code instructions to be executed to implement the steps. associated with the orbit control method 50 of the satellite 10. In a variant, the earth station comprises one or more programmable logic circuits, of the FPGA, PLD, etc. type, and / or specialized integrated circuits (ASIC) adapted to implement implement all or part of said associated steps of the orbit control method 50. In other words, the land station comprises a set of means configured in software (specific computer program product) and / or hardware (FPGA, PLD, ASIC, etc.) to implement the various stages of the program. orbit control method 50 which are executed by said earth station. More generally, it should be noted that the embodiments and embodiments considered above have been described by way of nonlimiting examples, and that other variants are therefore possible. In particular, the invention has been described by considering a maneuvering plane comprising two orbital control maneuvers per orbital period. Indeed, the invention makes it possible, in only two orbital control maneuvers per orbital period, to control in particular the inclination, the longitude and, in particular modes of implementation, the eccentricity of the orbit of the orbit. 10 while desaturating along three axes the device for storing the kinetic moment of said satellite 10. Nothing, however, excludes, according to other examples, having a different number of orbital control maneuvers per orbital period, higher or less than two. For example, the maneuvering plane may comprise an orbit control maneuver per orbital period, and may be defined, where appropriate, over a duration equal to or greater than two orbital periods.
    权利要求:
    Claims (19)
    [0001]
    CLAIMS1 - A method (50) for controlling the orbit of a satellite (10) in terrestrial orbit, in which the orbit of the satellite (10) is controlled by controlling a means of maneuvering the propulsion means (30, 31) , comprising at least one propellant, and means for displacing (20, 21) said propulsion means in a satellite reference centered on a center of mass of the satellite and comprising three axes X, Y and Z, the X axis being parallel to a speed vector of the satellite, the Z axis being directed towards the Earth, and the Y axis being orthogonal to the X and Z axes, characterized in that the displacement means being adapted to: - modify angles between a direction of thrusting each thruster and the axes X, Y respectively of the satellite reference, - moving each thruster, constant thrust direction in the satellite reference, so as to form a moment of any axis in a plane orthogonal to said thrust direction, the maneuvering plan comprises at least two orbit control maneuvers, thrust forces of the propulsion means during said two orbit control maneuvers being respective non-parallel thrust directions in an inertial frame, each of said thrust forces being determined so simultaneously controlling the inclination and longitude of the satellite orbit and forming a moment suitable for desaturating a device for storing the kinetic moment of said satellite in a plane orthogonal to the thrust direction of said thrust force.
    [0002]
    2 - Method (50) according to claim 1, characterized in that the maneuvering plane is determined under the constraint: IEN1 + EN2 + RN.sin (AT) I> r expression in which: - r is a strictly positive scalar value - EN1 is the ratio between an X-axis component and a Y-axis component of the thrust force of a first orbit control maneuver of the maneuvering plane, - EN2 is the ratio of a component along the X axis 3022530 30 and a component along the Y axis of the thrust force of a second orbit control maneuvering maneuvering plane, - RN corresponds to the ratio between a component along the Z axis and the component along the Y axis of the thrust force of the first or second maneuver control maneuver of the maneuvering plane, - AT is equal to 27. (T2 - T1 - Torb / 2) / Torb, expression where T1 and T2 are dates of the first and second orbit control maneuvers, and Torb is the orbital period of the satellite. 10
    [0003]
    3 - Method (50) according to one of the preceding claims, characterized in that the displacement means comprising an articulated arm (20, 21) carrying a propulsion means of propulsion, said articulated arm having at least three joints (22). , 23, 24) each having at least one rotational degree of freedom about an axis of rotation, the respective axes of rotation of adjacent joints being non-parallel for at least two pairs of adjacent hinges, the force thrust of said thruster is controlled by controlling the joints of the articulated arm.
    [0004]
    4 - Method (50) according to one of claims 1 to 2, characterized in that, the displacement means comprising a satellite attitude control device and an articulated arm carrying a propulsion means propulsion, said arm articulated having at least two joints each having at least one degree of freedom in rotation, the thrust force of said thruster is controlled by controlling the articulations of the articulated arm and the attitude of the satellite.
    [0005]
    5 - Process (50) according to one of claims 3 to 4, characterized in that, the articulated arm having at least one additional articulation, is controlled during at least one orbit control maneuver maneuvering plane , the propulsion means and the displacement means so as to form a thrust force determined so as to further control the eccentricity of the satellite orbit.
    [0006]
    6 - Method (50) according to one of claims 1 to 4, characterized in that the eccentricity of the orbit of the satellite is controlled by controlling 3022530 31 additional propulsion means (40) of said satellite, fixed orientation with respect to said satellite.
    [0007]
    7 - Method (50) according to one of the preceding claims, characterized in that the dates and / or the durations of said two control maneuvers 5 orbit of the maneuvering plane are determined so as to control the eccentricity of the orbit of the satellite during the maneuvering plan.
    [0008]
    8 - Method (50) according to one of the preceding claims, characterized in that an intermediate maneuvering plane, adapted to control only the orbit of the satellite, is determined by a terrestrial station and transmitted to the satellite, and the plane maneuver to implement is determined by the satellite according to said intermediate maneuvering plane.
    [0009]
    9 - Method (50) according to one of the preceding claims, characterized in that the maneuvering plane comprises at most two orbit control maneuvers per orbital period of the satellite. 15
    [0010]
    10 - Computer program product characterized in that it comprises a set of program code instructions which, when executed by a processor, implement a method of controlling the orbit of a satellite according to the one of claims 1 to 9.
    [0011]
    11 - Satellite (10) intended to be deployed in terrestrial orbit, comprising propulsion means (30, 31), comprising at least one propellant, and means (20, 21) for displacing said propulsion means in a a satellite reference center centered on a center of mass of said satellite and comprising three X, Y and Z axes, the X axis being parallel to a satellite speed vector, the Z axis being directed towards the Earth, and the Y axis being orthogonal To the X and Z axes, characterized in that the displacement means are adapted to: - modify angles between a thrust direction of each thruster and the axes respectively X, Y of the satellite reference, - move each thruster, direction of constant thrust 30 in the satellite reference, so as to form a moment of any axis in a plane orthogonal to said direction of thrust, and in that it comprises means adapted to control the propulsion means and the displacement means according to a a maneuvering plane 3022530 32 having at least two orbit control maneuvers, thrust forces of the propulsion means during said two orbit control maneuvers being respective non-parallel thrust directions in an inertial frame, each of said the thrust being determined so as to simultaneously control the inclination and longitude of the satellite orbit and to form a suitable moment to desaturate a device for storing the kinetic moment of said satellite in a plane orthogonal to the direction of thrust of the satellite. said pushing force.
    [0012]
    12 - Satellite (10) according to claim 11, characterized in that the displacement means 10 are arranged non-symmetrically with respect to the XZ plane formed by the X and Z axes of the satellite reference.
    [0013]
    13 - Satellite (10) according to one of claims 11 to 12, characterized in that the displacement means comprise an articulated arm (20, 21) carrying a propellant (30, 31) propulsion means, said articulated arm 15 having at least three joints (22, 23, 24) each having at least one degree of freedom in rotation about an axis of rotation, the respective axes of rotation of adjacent joints not being parallel to each other for at least two pairs of adjacent joints.
    [0014]
    14 - Satellite (10) according to claim 13, characterized in that the articulated arm (20, 21) comprises an additional articulation comprising at least one degree of freedom in rotation about an axis of rotation.
    [0015]
    15 - Satellite (10) according to one of claims 11 to 14, characterized in that it comprises additional propulsion means (40) fixed orientation relative to said satellite. 25
    [0016]
    16 - Satellite (10) according to one of claims 11 to 15, characterized in that the propulsion means (30, 31) carried by the displacement means (20, 21) are electric propulsion means.
    [0017]
    17 - orbit control system of a satellite (10) according to one of claims 11 to 16, characterized in that it comprises means 30 adapted to determine the maneuvering plane under constraint: IEN1 + EN2 + RN.sin (AT) I> r where: - r is a strictly positive scalar value, 3022530 33 - EN1 is the ratio between a component along the X axis and a component along the Y axis of the thrust force of a first maneuver control maneuver of the maneuvering plane, - EN2 corresponds to the ratio between a component along the X axis and a component along the Y axis of the pushing force of a second control maneuver. orbit of the maneuvering plane, - RN is the ratio between a component along the Z axis and the component along the Y axis of the thrust force of the first or second orbital control maneuver of plane 10 of maneuver, - AT is equal to 27. (T2 - T1 - Torb / 2) / Torb, expression in which T1 and T2 are dates of the first and second orbit control maneuvers, and Torb is the orbital period of the satellite.
    [0018]
    18 - System according to claim 17, characterized in that the means 15 adapted to determine the maneuvering plane are distributed between the satellite (10) and a ground station.
    [0019]
    19 - System according to claim 18, characterized in that an intermediate maneuvering plane, adapted to control only the orbit of the satellite, is determined by the earth station and transmitted to the satellite, and the maneuvering plane to implement is determined by the satellite as a function of said intermediate maneuvering plane.
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    同族专利:
    公开号 | 公开日
    FR3022530B1|2018-03-02|
    EP3157815A1|2017-04-26|
    CN106660641A|2017-05-10|
    US20170129627A1|2017-05-11|
    EP3381813B1|2021-05-05|
    EP3157815B1|2018-05-02|
    WO2015193499A1|2015-12-23|
    EP3381813A1|2018-10-03|
    US10232959B2|2019-03-19|
    ES2879228T3|2021-11-22|
    CN106660641B|2020-03-27|
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    优先权:
    申请号 | 申请日 | 专利标题
    FR1455630|2014-06-19|
    FR1455630A|FR3022530B1|2014-06-19|2014-06-19|METHOD FOR CONTROLLING THE ORBIT OF A SATELLITE IN TERRESTRIAL ORBIT, SATELLITE AND SYSTEM FOR CONTROLLING THE ORBIT OF SUCH A SATELLITE|FR1455630A| FR3022530B1|2014-06-19|2014-06-19|METHOD FOR CONTROLLING THE ORBIT OF A SATELLITE IN TERRESTRIAL ORBIT, SATELLITE AND SYSTEM FOR CONTROLLING THE ORBIT OF SUCH A SATELLITE|
    EP15732194.4A| EP3157815B1|2014-06-19|2015-06-19|Method for controlling the orbit of a satellite in earth orbit, satellite and system for controlling the orbit of such a satellite|
    EP18168546.2A| EP3381813B1|2014-06-19|2015-06-19|Satellite including electric propulsion means supported by moving means and additional electric propulsion means with fixed orientation|
    ES18168546T| ES2879228T3|2014-06-19|2015-06-19|Satellite comprising electric propulsion means supported by displacement means and additional electric propulsion means of fixed orientation|
    US15/318,670| US10232959B2|2014-06-19|2015-06-19|Method and system for controlling the orbit of a satellite in earth orbit|
    PCT/EP2015/063879| WO2015193499A1|2014-06-19|2015-06-19|Method for controlling the orbit of a satellite in earth orbit, satellite and system for controlling the orbit of such a satellite|
    CN201580037982.4A| CN106660641B|2014-06-19|2015-06-19|Method for controlling the orbit of a satellite in earth orbit, satellite and system for controlling the orbit of such a satellite|
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