• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
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    • 专利摘要:
    The invention relates to a satellite intended to be stationed in orbit around a celestial body, comprising a first steerable thrust direction electric thruster (21), a second steerable thrust direction electric thruster (22), and a an electric thruster (16) having a fixed orientation with respect to the satellite, a thrust line passing through the center of mass of the satellite. The satellite has two electric thruster power units (25, 26) and an electrical interconnection network (23) connecting a first power unit (26) to the first steerable thrust direction thruster (21) and to the fixed orientation thruster (16), and a second feed unit (25) to the second steerable thrust direction thruster (22) and the fixed orientation thruster (16), so that each of the units of power supply (25, 26) is capable of supplying either the associated steerable thrust steering thruster (22, 21) or the fixed orientation thruster (16).
    公开号:FR3032427A1
    申请号:FR1551034
    申请日:2015-02-10
    公开日:2016-08-12
    发明作者:Patrick Doubrere
    申请人:Airbus Defence and Space SAS；
    IPC主号:
    专利说明:

    [0001] The present invention is in the field of the positioning and post office maintenance of a spacecraft, more particularly a satellite, in its mission orbit around a celestial body. More particularly, the invention relates to a satellite intended to be stationed in a mission orbit around a celestial body, as well as a method of transferring such a satellite from an initial orbit into said mission orbit, and a method of controlling the orbit and attitude of such a satellite in said mission orbit. The invention finds a particularly advantageous, although in no way limiting, application in the case of telecommunications satellites intended to be placed in a geostationary orbit ("Geostationnary Orbit" or GEO) station equipped with electric-type propulsion means. Space vehicles such as artificial satellites are intended to be stationed in orbit around a celestial body, particularly in Earth orbit, especially in geostationary orbit, in order to carry out their mission, for example telecommunications, observation, etc. This posting is usually done in two stages. The first step is to launch the satellite in space, especially from the Earth's surface, by means of a specifically dedicated vehicle commonly called launcher vehicle, and to inject it into an initial orbit, known as an injection orbit.
    [0002] In the second stage, the satellite is transferred from this injection orbit to its mission orbit, also known as the final orbit. In a known manner, on its mission orbit, particularly in Earth orbit, a satellite 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.
    [0003] The orbit control consists of limiting the variations of the orbital parameters. 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, in terms of tilt, longitude and eccentricity, and is known also under the name of station keeping of the satellite ("station 5 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. Conventionally, several orbital control maneuvers are performed: - North / South maneuvers (N / S) make it possible to control the inclination of the satellite orbit, 15 - East-West maneuvers (E / O) to control the longitude of the satellite's orbit. Eccentricity can be controlled during E / O maneuvers or during N / S maneuvers. It is possible to define a satellite reference center centered on a center of mass of said satellite and having three axes X, Y and Z: when the satellite is stationed in its mission orbit, 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 maneuvers E / O requires X-axis thrust forces. In the present description, the satellite's center of mass is understood to mean the theoretical center of mass of the satellite; its actual center of mass may vary slightly over time depending on the amount of propellant in the tanks, the position / orientation of the payload equipment, etc.
    [0004] The satellites are conventionally equipped with propulsion means 3032427 3 able to perform on the one hand their transfer from the initial orbit to the mission orbit, and on the other hand their orbital maintenance in this orbit mission. These propulsion means may be of chemical type. Recently, electric propulsion, which provides better performance in comparison with chemical propulsion, has been used as a replacement for chemical propulsion to transfer and maintain orbit satellites. Satellites with electric propulsion means proposed by the prior art implement electric thrusters whose direction of thrust is steerable by mechanism. The orientation of the thrusters makes it possible in particular to control the position of the thrust direction relative to the center of mass of the satellite, and to pass from a configuration adapted to the electric transfer, in which the set of thrusters are oriented according to the same axis in the XZ plane (in general the Z axis of the satellite reference), to a configuration adapted to hold station.
    [0005] By way of example, US-A-5,443,231 discloses a satellite having four electric thrusters, each mounted on a mechanism for orienting its thrust direction. The transfer and station keeping of this satellite are performed by simultaneous or sequential implementation of two thrusters arranged diagonally, both in nominal mode in case of failure. Such a system, however, lacks robustness, the cases in which two thrusters on the same side of the satellite become defective being particularly detrimental to the mission of the satellite. In addition, the maneuvers necessary for controlling the orbit and attitude of such a satellite can prove to be complex to achieve, in particular as regards the control of eccentricity, and they are particularly likely to form moments can change the attitude of the satellite, which need to be corrected. The present invention aims to provide a satellite with electric propulsion means that allows to perform more simply and efficiently a large number of different maneuvers satellite orbit control and attitude.
    [0006] An additional object of the invention is that this satellite has a high degree of robustness in the case of failure of thrusters and / or power supply units and control thrusters, while having a mass and a limited cost.
    [0007] The invention also aims that the transfer phase of this satellite from the initial orbit to the mission orbit can be performed quickly. Thus, it is proposed according to the present invention a satellite intended to be stationed in a mission orbit around a celestial body, in particular a terrestrial orbit, in particular a geostationary orbit, which comprises, in a conventional manner in itself same, a so-called Earth face intended to be arranged facing the Earth when the satellite is stationary, and an opposite anti-Earth face, said satellite defining a satellite marker centered on a center of mass of the satellite and comprising three axes X, Y and Z , the Z axis being intended to be directed towards the Earth when the satellite is stationary, the X axis being parallel to a speed vector of the satellite, and the Y axis being orthogonal to the X and Z axes. This satellite comprises propulsion means comprising a first steerable thrust direction electric thruster, and a second steerable thrust direction electric thruster; at least two power supply units of an electric thruster, and an electrical interconnection network connecting a first power supply unit of an electric thruster to the first steerable thrust direction electric thruster, and a second drive unit of an electric thruster. supplying an electric thruster to the second steerable thrust direction electric thruster. A power unit, commonly referred to as PPU, for the English "Power Processing Unit", is defined in this description, in a conventional manner in itself, as an electronic unit which provides a main power supply adapted to the operation and operation of the power supply. control of a 3032427 5 electric thruster. The satellite according to the invention is furthermore such that the propulsion means also comprise an electric thruster with a fixed orientation relative to the satellite, a thrust line substantially aligned along the Z axis and passing through the center of mass of the satellite. . By substantially aligned means that the thrust line may be parallel to the Z axis (or be confused with this axis), that be inclined by a few degrees relative to the Z axis. The electrical interconnection circuit connects each of the first power unit of an electric thruster and the second power unit of an electric thruster to the fixed orientation electric thruster, such that each of said power units is capable of supplying either the associated steerable thrust steering electrical thruster or the fixed orientation electric thruster, depending on the particular needs of the mission.
    [0008] The fixed orientation electric thruster can advantageously be used for satellite orbit control and attitude maneuvers, both in the nominal operating mode and in the event of failure of one or more thrusters (s). ) electrical direction (s) directional thrust, and in particular for eccentricity control maneuvers. The moment applied to the satellite by this fixed orientation thruster is advantageously substantially zero as long as the actual center of mass of the satellite is close to its theoretical center of mass. In addition, in connection with this fixed orientation electric thruster, the particular configuration of the electrical interconnection network of the satellite 25 according to the invention, on the one hand, advantageously provides redundancy in the event of failure of the electric steering thruster of the steerable thrust, and, on the other hand, allows a greater diversity of maneuvers can be achieved by the propulsion means of the satellite, the power supplied by each power unit of an electric thruster can be allocated to 30 steerable thrust steering electrical thruster associated therewith, either to the fixed orientation electric thruster, as needed.
    [0009] Thus, in the event of failure of an electric steerable thrust direction thruster, the power unit associated with this inoperative thruster can be used to power the fixed thrust electric thruster. In particular, during the transfer phase of the satellite from the initial orbit 5 to the mission orbit, in the event of failure of an electric thruster of steerable thrust direction, the transfer can advantageously be continued by means of the thruster. fixed steering electric and steerable thrust steering electrical thruster remaining operational. For this purpose, the thrust of the latter is oriented substantially towards the center of mass of the satellite. The resulting total thrust is then greater than or equal to 90% of the nominal mode thrust. During the orbit control and attitude phase of the satellite, the fixed orientation electric thruster can also advantageously be used as a replacement for an electric steerable thrust direction thruster, in the event of a failure of the the latter, to ensure the control of eccentricity These advantageous results are obtained with only three electric thrusters in total, and two power units, when the prior art provides, to ensure the same degree of redundancy, four electric thrusters of steerable thrust direction, operating in pairs. The first steerable thrust direction electric thruster and the second steerable thrust direction electric thruster are preferably configured so as to be each capable of exerting a thrust comprising, in the satellite reference, a non-zero component along the Z axis and or a non-zero component along the Y axis. The thrust exerted by this first thruster and the thrust exerted by this second thruster are preferably in opposite directions along the Y axis. Preferably, the first thrust steering electric thruster. orientable and the second steerable directional thrust electric thruster are further disposed on distinct faces of the satellite, in particular arranged on opposite sides of the satellite, for example one on a conventionally called North face of the satellite, and the another on a face conventionally called South.
    [0010] According to particular embodiments, the invention furthermore satisfies the following characteristics, implemented separately or in each of their technically operative combinations. In general, apart from the particular features described above and hereinafter, in particular concerning its means of propulsion, the satellite is conventional in itself. It may include payloads, means for collecting solar energy, in the form of solar panels, generally deployable, communication means using deployable reflector antennas, and a service module, ensuring its basic functions, and including electrical power, control and navigation, telemetry and communication systems, etc., as well as the associated electrical wiring. It also has a reserve of propellant sufficient to ensure orbital transfer maneuvers and orbital maintenance, and if necessary for orbital change maneuvers for its transfer, end of life, in its orbit cemetery.
    [0011] All of these elements are well known to those skilled in the art and will not be described in detail in the present description. In particular embodiments of the invention, the fixed orientation electric thruster is disposed on the anti-Earth face of the satellite. It is therefore in particular no obstacle to the implementation and deployment of satellite payload instruments, which are conventionally installed on the satellite face of the satellite. In particularly advantageous embodiments of the invention, the electric thrusters of the propulsion means are all compatible with the same power supply units of an electric thruster.
    [0012] They are preferably identical.
    [0013] These electric thrusters may for example be Hall effect type, well known to those skilled in the art, it being understood that such an example is in no way limitative of the invention. In particular embodiments of the invention, the satellite 5 comprises at least three power supply units of an electric thruster and an electrical interconnection network connecting each of these power supply units of an electric thruster to at least an electric thruster means of propulsion of the satellite, so that can be implemented simultaneously the fixed orientation electric thruster and two thrusters electric steerable thrust direction. This third power supply unit of an electric thruster makes it possible to use, for the propulsion of the satellite, the simultaneous thrusts of both steerable thrust direction electric thrusters and fixed orientation electric thruster. During the phase of transfer of the satellite from the initial orbit to the mission orbit, this makes it possible to reduce the time required for the transfer, and thus to minimize in particular the exposure time of the satellite to the radiation of the Van Allen belts, as well as the cost of the mission. This is advantageously achieved by means of thrusters which are all of the same type, and use the same fuel, thus with a satellite equipment of limited cost and weight. This reduction in the duration of the orbit transfer phase is advantageously obtained by the implementation of material means (the fixed orientation electric thruster) which also advantageously constitute a redundancy solution vis-à-vis a propellant 25 steerable thrust steering electric. The satellite may comprise a number of power supply units of an electric thruster greater than three, for example equal to four or five. In particular embodiments of the invention, aimed at achieving one of the objectives that the present invention has set itself, which is to ensure a high degree of robustness of the satellite according to the invention in case of failure (s) , The satellite propulsion means comprise a first additional directional thrust direction electric thruster disposed on the same face of the satellite as the first steerable thrust direction electric thruster.
    [0014] Preferably, the satellite propulsion means further comprise a second electric thruster of additional steerable thrust direction disposed on the same face of the satellite as the second steerable thrust direction electric thruster. Thus, the satellite advantageously comprises two pairs of steerable thrust direction electric thrusters, preferably disposed respectively on two distinct faces of the satellite, preferably on the so-called north and south faces respectively of the satellite. A first pair of steerable thrust steering electric thrusters comprises the first steerable thrust steering electrical thruster and the first steerable steerable thrust steering electric thruster, and a second pair of steerable thrust steering electric thrusters comprises the second thruster steerable thrust steering electric and second steerable steerable thrust electric thruster.
    [0015] In such a configuration of the satellite according to the invention, quite advantageous in terms of redundancy in case of failure, in preferred embodiments of the invention, the satellite comprises a third power unit of an electric thruster. and an electrical interconnection network connecting this third power supply unit of an electric thruster to the first steerable additional directional thrust electric thruster and the second steerable additional thrust direction electric thruster. Thus, the electrical interconnection network preferably connects: a first power supply unit of an electric thruster to the first steerable thrust direction electric thruster and a fixed orientation electric thruster; a second thrust unit; feeding an electric thruster to the second steerable thrust direction electric thruster and the fixed thrust electric thruster, and a third thruster power unit to the first steerable thrust direction electric thruster and the second additional directional thrust steering electric thruster. Thus, with a limited number of electrical thrusters, and a number of power units of an electric thruster also limited, the satellite according to the invention has a high degree of redundancy, in case of failure of one or more electric thruster (s) and / or in the event of failure of one or more power unit (s) of an electric thruster. Compared to the configurations proposed by the prior art according to which the satellites are equipped with four steerable thrust electric thrusters, and two power supply units of an electric thruster, allowing the simultaneous operation of two electric thrusters, the combined implementation according to particular embodiments of the present invention, an additional electric thruster, fixed orientation with respect to the satellite and thrust line passing through the center of mass of the satellite and substantially aligned with the Z axis, and a third power unit of an electric thruster, allows both to reduce the duration of the orbit transfer phase to bring the satellite into its orbit mission (about a third ), And to provide a high degree of robustness, in case of failure of both an electric thruster and a power unit of a propellant. electrical power, both during the transfer phase and during the orbital holding phase of the satellite. In the case of a failure of a power unit, the duration of the transfer phase is simply reduced to the case of operation with two electric thrusters. This is achieved without significant overload of the power system, weight and cost of the satellite.
    [0016] The mission of the satellite is also robust to a double failure, including the simultaneous failure of an electric thruster directional thrust and a power unit of an electric thruster, or the simultaneous failure of two thrusters of the same pair of steerable thrust direction electric thrusters. In the latter case, the degradation of the mission time, due to less effective maneuvers, as explained below, is only slight. The satellite according to the invention may comprise a plurality of electric thrusters with a fixed orientation relative to the satellite, a thrust line substantially aligned along the Z axis and passing substantially through the center of mass of the satellite, arranged on the same face. of the satellite. In particular embodiments of the invention, the satellite comprises means for moving in the satellite reference of each of the steerable thrust direction electric thrusters of the satellite propulsion means. In particular embodiments of the invention, these displacement means are common for steerable thrust electric thrusters arranged on the same face of the satellite. The means of displacement in the satellite reference of each of the electric thrusters of steerable thrust direction can be formed by two-axis mechanisms, on each of which is individually mounted an electric propulsion thrust directional steering, each of these mechanisms allowing orientation of the associated thruster thrust line in a direction selected from the satellite coordinate system, generally substantially Z-axis, for the orbit transfer phase, and closest to the Y axis, aiming at approximately the center of mass of the satellite, for the orbital maintenance of the satellite. In particular embodiments of the invention, the displacement means of each of the steerable thrust direction electric thrusters comprise two articulated arms, each carrying a steerable thrust direction electric thruster and, if appropriate, the electric thruster. additional steerable thrust direction disposed on the same face of the satellite. Each of these articulated arms comprises at least three joints each having at least one degree of freedom in rotation about an axis of rotation. The thrust force of each thruster, in particular the direction of thrust and the point of application of the thrust force, is controlled by controlling the articulations of the articulated arm. In the context of such an embodiment of the invention, the fixed orientation electric thruster 10 according to the invention is particularly useful in the event of simultaneous failure of the two steerable thrust direction electric thrusters of the same pair, carried by the same articulated arm, or in case of failure of an articulated arm. Indeed, for maneuvers holding station, the fixed propeller then brings the component along the axis Z 15 missing, and the mission can continue with a single articulated arm. Similarly, during the transfer phase, the thrust can be effected by combining the thrust of the fixed orientation electric thruster and the thrust of the steerable thrust direction electric thruster remaining operational on the articulated arm. In this case, to maintain a torque-free thrust, the thrust of the steerable thrust steering electric thruster carried by the articulated arm is no longer aligned with that of the fixed thrust electric thruster. Typically, the angle between the outbreaks is less than 50 °, resulting in a loss of efficiency of less than 10%. According to another aspect, the present invention relates to a method for transferring a satellite according to the invention, corresponding to one or more of the above characteristics, of an initial orbit, in particular of an injection orbit. , in a mission orbit around a celestial body, notably a terrestrial orbit, and in particular a geostationary orbit. This method comprises a step of propelling the satellite by simultaneously using the fixed orientation electric thruster with respect to the satellite, Z-axis-aligned thrust line passing through the center of mass of the satellite, and 3032427 at least one steerable thrust steering electric thruster. More particularly, when the satellite comprises at least three power supply units of an electric thruster and an electrical interconnection network connecting each of these power supply units of an electric thruster to at least one electric thruster means of propulsion in such a way that both the fixed orientation electric thruster and two steerable thrust direction electric thrusters can be operated simultaneously, the transfer method preferably comprises a step of propelling the satellite by means of simultaneously the electric thruster of fixed orientation and at least two steerable thrust direction electric thrusters arranged on separate faces of the satellite. In particular embodiments of the invention, corresponding to a nominal mode of operation, for said propulsion stage, the thrust direction of each of the electric thrusters used is substantially aligned along the Z axis. The satellite mission orbit may for example be a terrestrial orbit, particularly the geostationary orbit. The initial orbit is then preferentially a low Earth orbit. The transfer of the satellite from the initial orbit into the mission orbit can be done in any conventional trajectory in itself, simple or sophisticated. According to another aspect, the present invention relates to a method of orbiting and attitude control of a satellite according to the invention, corresponding to one or more of the above characteristics, in a mission orbit around 25 a celestial body, especially a terrestrial orbit, in particular a geostationary orbit. This method comprises a step of propulsion of the satellite by the electric thruster of fixed orientation relative to the satellite. This step may be carried out separately, or simultaneously with a propulsion step by one or more steerable thrust direction electric thrusters. In particular, it makes it possible to provide eccentricity control 3032427. It can also be performed both in nominal mode, in case of failure. Another aspect of the invention resides in a remote control method of a satellite according to the invention, corresponding to one or more of the above characteristics, for the implementation of the steps of a method of transfer of the satellite according to the invention and / or steps of an orbit control method and attitude of the satellite according to the invention. According to this method, the satellite is remotely controlled by a control device, in particular on the ground, control signals being successively determined and sent to the satellite by this control device, for carrying out said steps. Another aspect of the invention relates to a control device which comprises means configured to control, preferably remotely, particularly from the earth's surface, a satellite according to the invention, responding to one or more of the above characteristics. by transmitting successive control signals to said satellite, to carry out the steps of the method according to the invention of transferring the satellite from the initial orbit into the mission orbit, then the steps of the method according to the invention of controlling the orbit and attitude of the satellite in the mission orbit. This control device, in particular ground control, is conventional in itself and may comprise one or more antennas for receiving signals from the satellite, and issue instruction signals towards the latter. It may include computers and means for processing and storing data received from the satellite. The latter is preferably preferably equipped with a control module, comprising in particular one or more processors, slaved to a communication module cooperating with the control device. The present invention also relates to a computer program product comprising a set of program code instructions which, when executed by a processor, implement a method of transferring a satellite according to the invention and / or a method for controlling the orbit and attitude of a satellite according to the invention.
    [0017] The invention will now be more specifically described in the context of preferred embodiments, which are in no way limiting, represented in FIGS. 1 to 6, in which: FIG. 1 schematically represents a satellite according to a Particular embodiment of the invention; FIG. 2 schematically represents a satellite according to a different embodiment of the invention; - Figure 3 schematically shows a satellite according to a preferred embodiment of the invention; FIG. 4 shows a partial view of a satellite according to one particular embodiment of the invention, illustrating the operation of the fixed orientation electric thruster; FIG. 5a shows a partial view of a satellite according to one particular embodiment of the invention, in an orbit transfer configuration for the positioning of the satellite on its mission orbit; FIG. 5b shows a partial view of the satellite of FIG. 5a, in the configuration of North / South maneuvers; and FIG. 6 shows a diagram illustrating the electrical interconnection network of a satellite according to a particular embodiment of the invention. The invention will be described hereinafter with reference to the particular nonlimiting example of a satellite 10 to be placed in position in geostationary orbit. Nothing, however, excludes, according to other examples, considering 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. A satellite 10 according to a particular embodiment of the invention is shown schematically in FIG.
    [0018] This satellite comprises, in a conventional manner in itself, a so-called Earth face 101, intended to be directed towards the Earth when the satellite is stationary, and an opposite anti-Earth face 102. The Earth face 101 generally carries the satellite payload instruments. The satellite 10 5 also has a so-called south face 103 and an opposite north face 104 opposite. The satellite 10 defines a satellite reference having three X, Y and Z axes. More particularly, the X axis is parallel to a speed vector of the satellite 10 in an inertial frame, the Z axis is directed towards the center of the Earth when the satellite is in geostationary orbit, and the Y axis is orthogonal to the axes X and X. As illustrated in Figure 1, the satellite 10 comprises for example a body 20, and two solar generators 11, 11 'from The two solar generators 11, 11 'are, for example, rotatably mounted relative to the body 20 of the satellite 10, around a same axis of rotation. A first solar generator 11 deploys from the south face 103 of the satellite 10, and a second solar generator 11 'deploys from its north face 104. The satellite 10 further comprises at least two thrust direction electric thrusters, carried for one by the south face 103 of the satellite 10, and for the second pair by its north face 104. The satellite 10 preferably comprises at least two pairs of steerable thrust direction electric thrusters, carried for the first pair by the South face 103 of the satellite 10, and for the second pair by its north face 104. These thrusters are not shown in Figure 1 but the associated thrust direction is illustrated at 12, 12 'for the pair of thrusters carried by the South face 103 of the satellite 10, and at 13, 13 'for the pair of thrusters carried by its north face 104. In each pair, a so-called nominal thruster is generally implemented in operation no minal, and a second thruster, said additional thruster, provides redundancy for cases of failure of the nominal thruster. Each pair of thrusters is carried by an articulated arm 14, 15, 3032427 17 each of these arms 14, 15 having three joints each having at least one degree of freedom in rotation relative to an axis of rotation. An exemplary embodiment of such arms will be described in more detail in the present description.
    [0019] The satellite 10 may comprise one or more additional directional thrust direction electric thrusters. The satellite 10 further comprises a thruster of fixed orientation 16, whose thrust line, shown at 17 in FIG. 1, is substantially aligned along the axis Z, which is the axis intended to be directed towards the Earth 10 when the satellite 10 is stationed in the geostationary orbit, and passes through the center of mass of the satellite 10 (not shown in this figure). It may further comprise one or more electrical thrusters of fixed orientation relative to the additional satellite. The set of electric thrusters, whether of steerable thrust direction or of fixed orientation with respect to the satellite 10, are preferably, but not necessarily, identical, so that they can be powered and controlled by the same units. power supply. These electric thrusters may for example be of the Hall effect type and have a power of between 2.5 and 5 kW each.
    [0020] The satellite 10 comprises, in a conventional manner in itself, a propellant reservoir, not shown in the figures, adapted to receive a volume of propellants in the form of gas, for example xenon, for the supply of propellants. electric. A satellite variant 10 according to the invention is shown schematically in FIG. 2. This satellite 10 is identical to that described above with reference to FIG. 1, with the difference that the steerable thrust direction electric thrusters are carried. each by an individual articulated arm 18, 18 'for the thrusters carried by the south face 103 of the satellite 10, and 19, 19' for the thrusters carried by the north face 104 of the satellite 10, 30 these arms each having two joints each comprising at least 3032427 18 a degree of freedom with respect to an axis of rotation. These articulated arms 18, 18 'and 19, 19' each make it possible to control the orientation of the direction of thrust of the steerable thrust direction electrical thruster which is associated with it, in particular in the XZ plane for a satellite transfer phase. 10 from an initial orbit to its geostationary mission orbit, and substantially along the Y axis for station keeping maneuvers in that orbit. Figure 3 shows a more detailed representation of a satellite according to a particularly advantageous embodiment of the invention.
    [0021] For the purposes of the description of this figure, the satellite 10 is associated with a satellite reference centered on a center of mass O of the satellite 10 and comprising three axes X, Y, Z. More particularly, the axis X is parallel to a satellite speed vector 10 in inertial reference, the Z axis is directed towards the center of the Earth, 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, and the unit vector uy is oriented so that the together (ux, uy, uz) constitutes a direct orthonormal basis of the satellite reference. In the remainder of the description, one places oneself in a nonlimiting manner in the case where the body 20 of the satellite 10 is substantially in the form of a rectangular parallelepiped. The body 20 thus comprises six faces two by two parallel.
    [0022] Furthermore, in the case where the attitude of the satellite 10 is controlled, for the purposes of the mission of said satellite 10, is placed in a non-limiting manner so as to be placed in a set-point attitude, called "mission attitude". in which: a face of the body 20 of the satellite 10, called the Earth face 101, carrying for example an instrument with a payload of said satellite 10, is directed towards the Earth and is substantially orthogonal to the Z axis; ; the opposite face to the face Earth, then arranged on the opposite side to the Earth, is said anti-Earth face 102; the two opposite faces of the body 20 of the satellite 10 on which are arranged the two solar generators 12, respectively said South face 103 and North face 104, are substantially orthogonal to the Y axis; and the last two opposite faces of the body of the satellite 10, respectively 105 and 106, are substantially orthogonal to the axis X. The satellite 10 also comprises a set of actuators adapted to control the orbit and the attitude of said satellite 10, and a control device 10 (not shown in the figures) of said actuators. For the purposes of attitude control, the satellite 10 preferably comprises a kinetic moment storage device. The satellite 10 further comprises propulsion means, comprising a set of electric thrusters, more particularly: an electric thruster 16 of fixed orientation with respect to the satellite 10, disposed on the anti-Earth face 102 of the satellite 10, and whose thrust line is substantially aligned along the Z axis and passes through the center of mass O of the satellite; and two pairs of steerable thrust direction electrical thrusters: a first pair of electric thrusters 21, 21 'is carried by the south face 103 of the satellite 10, and a second pair of electric thrusters 22, 22' is carried by the North face 104 of the satellite. The point of attachment of the fixed orientation thruster 16 to the anti-Earth face 102 of the satellite 10 corresponds substantially to the orthogonal projection 25 of the theoretical center of mass of the satellite 10 on said anti-Earth face 102. Thus, the moment applied to the satellite 10 by the fixed orientation thruster 16 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 16 orientation 3032 42 7 20 fixed relative to the satellite 10. The fixed orientation thruster 16 can be implemented to ensure the transfer satellite 10 of the initial orbit into its geostationary mission orbit, or, at station, to control the eccentricity of the orbit. It can be activated simultaneously with the electric thrusters 21, 21 ', 22, 22', and / or during dedicated eccentricity control maneuvers, distinct from the N / S and E / O control maneuvers of the orbit. Each pair of steerable thrust direction electrical thrusters 21, 21 'and 22, 22' is associated with means for moving said thrusters 10 in the satellite fixture, adapted to simultaneously control the inclination and longitude of the satellite orbit. 10. More particularly, these displacement means are suitable for: modifying angles between a thrust direction of each thruster and the X, Y axes respectively of the satellite marker, moving each thruster, with constant thrust direction, in the reference mark satellite, 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 a satellite control device 10, by controlling the propulsion means and the displacement means according to a maneuvering plane comprising orbit control maneuvers during from which the propulsion means are activated. With such means of displacement, it is understood that it is possible, during a single 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.
    [0023] In the example illustrated in FIG. 3, the displacement means 3032427 21 comprise two articulated arms 14, 15, each articulated arm 14, 15 carrying two electric thrusters 21, 21 'and 22, 22'. In the nonlimiting example illustrated in FIG. 3, the articulated arms 14, 15 are arranged on the south face 103 and the north face 104 respectively of the body 20 of the satellite 10. The articulated arms 14, 15 are for example respectively for the South control and the North control of the inclination of the orbit of the satellite 10, alternately activating either a thruster 21, 21 'or a thruster 22, 22'. The articulated arm 14 is preferably fixed to the south face 103 at a fixed point which is offset, along the Z axis, relative to the orthogonal projection of the theoretical center of mass O of the satellite 10 on said south face 103. analogously, the articulated arm 15 is preferably fixed to the north face 104 at a fixed point which is offset along the Z axis relative to the orthogonal projection of the theoretical center of mass O of the satellite 10 on said North face 15 104. In this configuration, the thrust force of the steerable thrust direction electric thrusters respectively 21, 21 ', and 22, 22' comprises in north / south control a component along the Z axis without forming a moment. Such a configuration is however not limited to the invention.
    [0024] Each of the two articulated arms 14, 15 comprises at least three joints 141, 142, 143 and 151, 152, 153, respectively. Each of these articulations comprises at least one degree of freedom in rotation about an axis of rotation. The joints 141 and 142, and 151 and 152, are interconnected by a connection respectively 144 and 154, while the joints 142 and 143, and 152 and 153, are interconnected by a connection 145 and 155, respectively. platinum 146, 156 extends from the terminal joint 143, 153 of each articulated arm 14, 15 and carries the electric thrusters 21, 21 'and 22, 22'.
    [0025] Each articulated arm 14, 15 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 electric thrusters that it carries. . For this purpose, the control device controls the angles of rotation of the joints 141, 151, 142, 152, and 143, 153 designated respectively by 01, 02 and 03. The various electric thrusters of the satellite 10 according to the invention can be implemented both for the satellite transfer phase of an initial orbit, in particular an injection orbit into which it has been injected by a launcher vehicle, into its geostationary mission orbit, as for the control of orbit and satellite attitude to post. For all of these phases, the fixed orientation electric thruster 16 with respect to the satellite 10 has a thrust of fixed direction, substantially aligned along the Z axis of the satellite, as indicated schematically in FIG. 4.
    [0026] This fixed orientation electric thruster 16 may be implemented during the orbit transfer phase, in nominal operation, to reduce the transfer time, or in the event of failure of one or more steering thrusters. steerable thrust 21, 21 ', 22, 22'. It can also be implemented for the orbit and attitude control of the satellite 10 stationed in the geostationary orbit, in particular for the eccentricity control maneuvers, in nominal operation, or in case of failure, in particularly in the event of failure of a pair of steerable thrust direction electric thrusters 21, 21 'or 22, 22' or of one of the articulated arms 14, 15. The fixed orientation electric thruster 16 thus allows 25 compensate for the thrust along the Z axis created by the steerable thrust direction electric thrusters that remain operational. The steerable thrust direction electrical thrusters 21, 21 ', 22, 22' can be implemented for both the orbit transfer phase and the orbit control and attitude phase of the satellite 10. Indeed, they are moved in the satellite reference, so as to orient their direction of thrust appropriately. FIGS. 5a and 5b illustrate examples of positioning of the steerable thrust direction electrical thrusters 21, 21 'carried by the articulated arm 14. For the orbit transfer phase, or for certain orbital control and control maneuvers. attitude of the satellite 10, as shown in Figure 5a, the electric thrusters 21, 21 'are placed in a configuration in which they are able to exert a thrust along the Z axis. For some control maneuvers of orbit and attitude of the satellite 10, as illustrated in FIG. 5b, the electric thrusters 21, 21 'are placed in a configuration in which they are able to exert thrust in a different direction in the satellite reference. The satellite 10 further comprises at least two power supply units of an electric thruster, referred to as PPU in the remainder of this description, which are conventional in themselves. Preferably, it comprises at least three PPUs 24, 25, 26, and an electrical interconnection network 23, an exemplary embodiment of which is illustrated in FIG. 6, connecting each of these PPUs to one or more electric thrusters of the satellite. 10. In the particularly advantageous embodiment illustrated in FIG. 6, the electrical interconnection network 23 interconnects the following different components: a first power supply unit of an electric thruster 26, called PPU1, is connected on the one hand to the nominal thruster 21 of the first pair of steerable thrust direction electrical thrusters, said S1, and on the other hand to the fixed orientation electric thruster 16; a switch, not shown in the figure, allows the power supply unit PPU1 26 to supply either the electric thruster S1 21 or the fixed orientation electric thruster 16; a second power supply unit of an electric thruster 25, called PPU2, is connected on the one hand to the nominal thruster 22 of the second pair of steerable thrust direction electric thrusters, called N1, and on the other hand 3032427 part of fixed orientation electric thruster 16; a switch, not shown in the figure, allows the power supply unit PPU2 25 to supply, either, the electric thruster N1 22 or the fixed orientation electric thruster 16; 5 - and a third power supply unit of an electric thruster 24, called PPU3, is connected on the one hand to the additional thruster 21 'of the first pair of steerable thrust steering electric thrusters, said S2, and other part of the additional thruster 22 'of the second pair of steerable thrust direction electric thrusters, said N2; a switch, not shown in the figure, allows the power supply unit PPU3 24 to supply either the electric thruster S2 21 'or the electric thruster N2 22'. The connection between the PPUs and the electric thrusters is carried out by conventional electrical wiring in itself.
    [0027] A relay box 27, also conventional in itself, is arranged upstream of the fixed orientation electric thruster 16, on the feed paths from the PPU1 supply unit 26 and the PPU2 supply unit. 25, to allow the supply of the fixed orientation electrical thruster 16 either from the first PPU1 26 or from the second PPU2 25, at choice. The satellite 10 having the above characteristics advantageously has a particularly high degree of robustness, and can easily adapt to a large number of various faults, including double faults, so that its operation, and in particular the duration of 25 its mission, is little or not impacted by failures, and this for a difference in mass and cost compared to satellites of the relatively low prior art. This advantage is added to that of reducing the duration of the orbit transfer phase, from the injection orbit to the geostationary orbit, which makes it possible in particular to minimize the exposure time of the satellite 10 to 30. radiation from Van Allen's belts.
    [0028] For the implementation of the orbit transfer phases and the orbit control and attitude control phases, the satellite 10 can be controlled remotely by a control device, particularly on the ground, in a conventional manner in itself. -even.
    [0029] This remote control device is configured to control the different phases implemented by the satellite 10. For this purpose, the control device and the satellite 10 each comprise conventional means of remote communication. The control device is adapted to determine control signals which are sent to the satellite 10. These control signals are for example determined as a function of measurement signals received from the satellite 10, which are determined by different sensors (gyroscope, stellar sensor , etc.) of the latter. The satellite 10 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 be executed to implement the different steps of a satellite control method 10. In a variant, the control device also comprises one or more programmable logic circuits, of FPGA, PLD, etc. type, and / or specialized integrated circuits (ASIC) adapted to put all or some of said steps of the control method. In other words, the control device comprises a set of means configured in a software (specific computer program product) and / or hardware (FPGA, PLD, ASIC, etc.) way to implement the various steps of FIG. a method of transferring the satellite 10 from the injection orbit into the mission orbit, and then to a method of orbiting and attitude control of the satellite 10 in the mission orbit.
    权利要求:
    Claims (16)
    [0001]
    REVENDICATIONS1. Satellite (10) intended to be stationed in a mission orbit around a celestial body, having a so-called Earth face (101) intended to be disposed facing the Earth when said satellite (10) is stationary, and a opposite anti-Earth face (102), said satellite (10) defining a satellite marker centered on a center of mass of the satellite and having three axes X, Y and Z, the Z axis being intended to be directed towards the Earth when said satellite (10) is stationed, said satellite (10) comprising: - propulsion means comprising a first steerable thrust direction electric thruster (21), and a second steerable thrust direction electric thruster (22), - at at least two power supply units of an electric thruster (25, 26) and an electrical interconnection network (23) connecting a first power supply unit of an electric thruster (26) to said first thrust steering electrical thruster e (21), and a second power supply unit of an electric thruster (25) to said second steerable thrust electric thruster (22), characterized in that said propulsion means further comprises an electric thruster (16) a fixed orientation with respect to the satellite (10), a line of thrust aligned along the Z axis and passing through the center of mass of the satellite (10), and in that said electrical interconnection network (23) connects each said first power supply unit of an electric thruster (26) and said second electric thruster power unit (25) to said fixed orientation electric thruster (16), such that each of said power units The power supply (25, 26) is capable of supplying either the associated steerable thrust steering electric thruster (22, 21) or said fixed orientation electric thruster (16). 3032427 27
    [0002]
    The satellite (10) of claim 1, wherein the fixed orientation electric thruster (16) is disposed on the anti-Earth face (102) of the satellite (10).
    [0003]
    The satellite (10) according to any one of claims 1 to 2, comprising at least three power supply units of an electric thruster (24, 25, 26) and an electrical interconnection network (23) connecting each said power supply units of an electric thruster to at least one electric thruster of said propulsion means so that said fixed orientation electric thruster (16) and two steerable thrust power thrusters can be simultaneously operated; .
    [0004]
    4. Satellite (10) according to any one of claims 1 to 3, wherein the propulsion means comprise a first electric thruster of additional steerable thrust direction (21 ') disposed on the same face (103) of said satellite as said first steerable thrust direction electric thruster (21).
    [0005]
    5. Satellite (10) according to claim 4, wherein the propulsion means comprise a second electric thruster of additional steerable thrust direction (22 ') disposed on the same face (104) of said satellite as said second electric thruster direction of steerable thrust (22).
    [0006]
    The satellite (10) of claim 5 including a third power supply unit of an electric booster (24) and an electrical interconnection network connecting said third power supply unit of an electric booster (24) to said first power supply unit (24). additional steerable thrust steering electric thruster (21 ') and said second steerable steerable thrust steering electric thruster (22').
    [0007]
    7. Satellite (10) according to any one of claims 1 to 6, comprising a plurality of electrical thrusters (16) of fixed orientation relative to the satellite (10), Z-aligned thrust line and passing 3032427 28 substantially by the center of mass of the satellite, arranged on the same face (102) of said satellite (10).
    [0008]
    8. Satellite (10) according to any one of claims 1 to 7, comprising means for moving each of the electric thrusters directional steerable thrust (21, 21 ', 22, 22') means of propulsion in the satellite reference.
    [0009]
    9. Satellite (10) according to claim 8, wherein the displacement means are common for steerable thrust direction electric thrusters arranged on the same face of said satellite. 10
    [0010]
    10. Satellite (10) according to any one of claims 8 to 9, wherein the means for moving each of said steerable thrust direction electric thrusters (21, 21 ', 22, 22') comprise two articulated arms (14 15) each carrying a steerable thrust direction electric thruster (21, 22), and if desired the steerable thrust direction electric thruster disposed on the same face of said satellite (21 ', 22'), each of said arms articulated (14, 15) having at least three joints (141, 142, 143 and 151, 152, 153) each having at least one degree of freedom in rotation about an axis of rotation.
    [0011]
    The satellite (10) according to any one of claims 1 to 10, wherein the electric thrusters (16, 21, 21 ', 22, 22') of the propulsion means are all compatible with the same power units. an electric thruster, and preferably identical.
    [0012]
    A method of transferring a satellite (10) according to any one of claims 1 to 11 from an initial orbit into a mission orbit of said satellite (10) around a celestial body, characterized in that it comprises a step of propelling said satellite (10) simultaneously by means of the fixed orientation electric thruster (16) relative to the satellite (10) and at least one steerable thrust direction electric thruster (21, 21 ', 22, 22 '). 3032427 29
    [0013]
    13. Transfer method according to claim 12, wherein the satellite (10) comprising at least three power supply units of an electric thruster (24, 25, 26) and an electrical interconnection network (23) connecting each said power supply units of an electric thruster to at least one electric thruster of said propulsion means so that said fixed orientation electric thruster (16) and two steerable thrust power thrusters can be simultaneously operated; said transfer method comprises a step of propelling said satellite (10) by means of simultaneously said fixed orientation electric thruster (16), and at least two steerable thrust direction electric thrusters (21, 21 ', 22). , 22 ') arranged on separate faces (103, 104) of said satellite.
    [0014]
    14. The orbit control and attitude method of a satellite (10) according to any one of claims 1 to 11, in a mission orbit around a celestial body, characterized in that it comprises a step of propelling said satellite (10) by the electric thruster (16) fixed orientation relative to said satellite (10).
    [0015]
    15. A method of remote control of a satellite (10) for carrying out the steps of a transfer method according to any one of claims 12 to 13 and / or steps of a control method of orbit and attitude according to claim 14, wherein said satellite (10) is controlled remotely by a control device, control signals being successively determined and sent to said satellite (10) by said control device for the realization said steps. 25
    [0016]
    16. Computer program product characterized in that it comprises a set of program code instructions which, when executed by a processor, implement a method of transferring a satellite (10) according to the present invention. any one of claims 12 to 13 and / or an orbit control and attitude method of a satellite (10) according to claim 14.
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    同族专利:
    公开号 | 公开日
    EP3201090A1|2017-08-09|
    US9926087B2|2018-03-27|
    FR3032427B1|2017-03-10|
    ES2661024T3|2018-03-27|
    WO2016128389A1|2016-08-18|
    EP3201090B1|2017-12-13|
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    法律状态:
    2016-02-26| PLFP| Fee payment|Year of fee payment: 2 |
    2016-08-12| PLSC| Publication of the preliminary search report|Effective date: 20160812 |
    2017-02-28| PLFP| Fee payment|Year of fee payment: 3 |
    2018-02-28| PLFP| Fee payment|Year of fee payment: 4 |
    2020-02-28| PLFP| Fee payment|Year of fee payment: 6 |
    2021-02-26| PLFP| Fee payment|Year of fee payment: 7 |
    2022-02-24| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1551034A|FR3032427B1|2015-02-10|2015-02-10|SATELLITE HAVING ELECTRICAL PROPULSION MEANS, METHOD OF POSTING SUCH A SATELLITE AND METHOD FOR MAINTAINING SATELLITE|FR1551034A| FR3032427B1|2015-02-10|2015-02-10|SATELLITE HAVING ELECTRICAL PROPULSION MEANS, METHOD OF POSTING SUCH A SATELLITE AND METHOD FOR MAINTAINING SATELLITE|
    EP16705070.7A| EP3201090B1|2015-02-10|2016-02-09|Satellite comprising electrical propulsion means, method for placing such a satellite in a station and method for keeping said satellite in its station|
    ES16705070.7T| ES2661024T3|2015-02-10|2016-02-09|Satellite with electric propulsion means, procedure for setting up such satellite and maintenance procedure in position of said satellite|
    PCT/EP2016/052710| WO2016128389A1|2015-02-10|2016-02-09|Satellite comprising electrical propulsion means, method for placing such a satellite in a station and method for keeping said satellite in its station|
    US15/549,165| US9926087B2|2015-02-10|2016-02-09|Satellite comprising electrical propulsion means, method for placing such a satellite in a station and method for keeping said satellite in its station|
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