• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to a method for controlling attitude control of a satellite with respect to an orbital reference system comprising a speed axis, an orbital axis and a Nadir axis; the satellite moving in the direction of the speed axis, the satellite comprising an optical instrument having a target axis, a solar generator defining a functional surface having a normal, an attitude control device and a control unit. The method includes a guide control transmission step (104) for pointing the line of sight of the optical instrument toward areas to be imaged or for directing the normal to the functional surface in the direction of solar radiation. The guidance commands are only commands in rotation of the satellite around the speed axis, the rotation angle around the orbit axis and the Nadir axis in the orbital frame being kept substantially zero.
    公开号:FR3047813A1
    申请号:FR1651237
    申请日:2016-02-16
    公开日:2017-08-18
    发明作者:Emmanuel Giraud
    申请人:Airbus Defence and Space SAS；
    IPC主号:
    专利说明:

    The invention relates to the spatial domain, and more particularly, the invention relates to guiding the attitude of a satellite, satellite, pluralities of satellites and associated computer program. several optical satellites orbiting a star, especially the Earth.
    An optical satellite is a satellite that includes an optical instrument and whose main mission is related to this instrument. This is for example a satellite whose mission is the shooting of the surface of the Earth, or any other star, for monitoring or mapping purposes.
    An optical instrument for space missions is typically formed of at least one dioptric, catadioptric or mirror lens having an optical axis, and the primary mirror for focusing the light rays, to obtain an image in a focal plane equipped with imaging systems. detection. The line of sight can be confused with the optical axis of the lens or can form an angle with it by means of deflection mirrors. When the optical instrument is a shooting instrument, that is to say comprising at least one sensor making it possible to form an image of a region, for example a region of the terrestrial ground, the optical instrument defines also a field of view corresponding to the truncated cone extending from the functional surface of the sensor, that is to say the surface of the sensor on which the images are formed to the region of shooting.
    The satellite furthermore includes so-called secondary equipment, contributing to the overall good functioning of the satellite, such as propellers for correcting the trajectory of the satellite, reaction wheels for modifying the angular orientation of the satellite, radiative panels for evacuating heat or again a solar generator or a battery for receiving and managing power. The orientation of the line of sight towards the area to be observed can be done by a mirror of mobile referral in front of the telescope, the satellite keeping a fixed orientation, as for the satellites SPOT 1 to 5. In these satellites, one generally uses an orientable solar generator pointing towards the sun. A disadvantage of these satellites according to the state of the art is that the presence of the moving mirror in relation to the satellite body decreases the quality of the shooting. Indeed, the introduction of mobile elements is detrimental to the stability of the line of sight of the optical instrument. In addition, the presence of means to obtain the displacement of the mobile deflection mirror increases the mass of the satellite and its size, increasing the costs of design and launch of the satellite.
    In more recent satellites such as PLEIADES satellites, the satellite is oriented so that the telescope's aiming axis is aimed at the area to be observed. In addition, the solar generator is fixed. We then speak of the attitude, or angular attitude, of the satellite as being the angular position of a frame linked to the satellite with respect to an external frame of reference, for example linked to the orbit of the satellite.
    These satellites can operate according to two statuses, namely an operating status in which the optical instrument is implemented to take a shot, in practice a succession of shots, and a waiting status in which the optical instrument is not used. This waiting status recharges the satellite batteries. In this waiting status, the attitude of the satellite is such that the planar solar panels are normal to the direction of the sun. PLEIADES satellites operate on this basis.
    Between one status and another, the attitude of the satellite and the orientation of its equipment are modified in the external reference system so as to optimize either the shooting or the recharging of the batteries by orienting the generator perpendicular to the sun's rays. In the case of a LEO (Low Earth Orbit) orbit, in which the satellite is at an altitude such that drag is important, this change of orientation between the two operating states poses the problem of optimizing the surfaces. to reduce drag. Indeed, to reduce the drag, the frontal surface of the satellite following the direction of its displacement on its orbit must be reduced. However, the smaller the surface, the less it is available to attach equipment. The two operating states imply that the front face of the satellite changes, so that several faces of the satellite must have their surface reduced, which in practice reduces the possibility of reducing the drag of the satellite by changing its shape.
    This type of satellite requires inertial actuators (typically CMG) of attitude performing around the three axes X, Y, Z, according to Figure 1, to quickly orient the satellite to the target area. These CMG actuators are expensive and heavy.
    There is therefore a need for a new method to overcome in particular the disadvantages of the state of the art mentioned above. To this end, the subject of the invention is a method for controlling the attitude of a satellite with respect to an orthogonal orbital reference system comprising a so-called speed axis, an orbital axis and an Nadir axis, along a a portion of orbit around the Earth, said portion of orbit being illuminated by solar radiation; the satellite moving in the direction of the speed axis, the satellite comprising a main body, an optical instrument having a fixed axis of view relative to the main body, at least one solar generator fixed relative to the main body defining a functional surface whose normal has at least one component perpendicular to the speed axis, at least one attitude control device and a control unit connected to the attitude control device, said method comprising a control guide transmission step of the control unit at the attitude control unit for pointing the line of sight of the optical instrument towards regions to be imaged or for directing the normal to the functional surface in the direction of solar radiation, characterized in that that the guidance commands are only commands in rotation of the satellite around the axis speed, the angle of rotation around of the axis of orbit and the axis Nadir in the orbital reference being maintained substantially harmful.
    According to particular embodiments, the artificial satellite comprises one or more of the following characteristics: the guidance commands comprise commands capable of rotating the satellite around the speed axis over angular ranges to scan with the axis of sighted a portion of the earth; the method described above is implemented by several satellites; the normal to the functional surface of the solar generator is parallel to the line of sight of the optical instrument and oriented in the opposite direction; the angles of rotation about the speed axis are limited to a predetermined clamping angle, said guide angle being defined with respect to the orbital axis; the clamping angle is 50 °; when the illumination of the satellite is zero, for example during an eclipse, the satellite is rotated about the speed axis so as to point the line of sight towards the Earth; and the line of sight is perpendicular to the speed axis. The subject of the invention is also a satellite comprising a main body, an optical instrument whose axis of view is fixed with respect to the main body, at least one solar generator fixed with respect to the main body, at least one control device. attitude and a control unit connected to the attitude control device, the control unit being adapted to execute the guidance control method according to any one of the preceding claims, the attitude control device is suitable for rotating the satellite about a first axis, a second axis and a third axis, said first, second axis and third axis being perpendicular to each other, said third axis being parallel to the line of sight of the optical instrument and oriented in the same direction, characterized in that the torque capacity of the attitude control device according to the second axis and / or according to the third sth axis is less than 40% of the torque capacity along the first axis.
    According to particular embodiments, the artificial satellite comprises one or more of the following features: - which comprises an interface device intended to cooperate with a complementary interface device of a launcher or of a satellite, and comprising an intermediate structure connecting the body of the satellite to the interface device, the line of sight of the optical instrument being oriented towards the interface device; and the normal to the functional surface of the solar generator is parallel to the line of sight of the optical instrument and oriented in the opposite direction. The invention also relates to a plurality of satellites shaped according to the characteristics mentioned above. The plurality of satellites is intended to operate in a constellation, said plurality of satellites being able to be guided in orbit by the method mentioned above.
    Finally, the 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 the method mentioned above. Other characteristics and advantages of the invention will emerge in the light of the description of embodiments accompanied by the figures in which:
    Figure 1 is a schematic representation of a satellite orbiting the Earth and its local orbital repository.
    Figure 2A is a diagram showing the steps of the method according to the invention.
    FIG. 2B is a diagram illustrating an exemplary rotation of a satellite guided by the method according to the invention.
    Figure 3 is a schematic representation of three satellites of a constellation guided by the method according to the invention in a shooting status.
    Figures 4 and 5 are diagrams illustrating examples of shooting by a satellite (Figure 4) and two satellites (Figure 5) guided by the method according to the invention.
    Figure 6 is a schematic representation in the XZ plane of a satellite in a wait status.
    Figure 7 is a schematic plan view of the satellite of Figure 6 in the YZ plane.
    FIGS. 8 to 10 each represent an example of positioning a solar generator on a satellite guided by the method according to the invention.
    FIG. 11 is a graph illustrating the attitude of a satellite according to a first positioning of the solar generator guided by the method according to the invention in a waiting status.
    FIG. 12 is a graph illustrating the attitude of a satellite according to a second positioning of the solar generator guided by the method according to the invention in a waiting status.
    Figure 13 is a schematic three-dimensional view of another example of a satellite design guided by the method according to the invention.
    In FIG. 1, there is shown an example of a shooting satellite 1 comprising an optical instrument 2, moving along an orbit A around a star such as Earth T.
    Conventionally, the satellite 1 has been previously placed in a launcher, through which the satellite 1 is launched into space. The satellite 1 is dropped by the launcher into space. Then he rejoins his planned operational orbit A.
    According to the example in FIG. 1, but in a nonlimiting manner, the satellite comprises a parallelepiped body 3 of center of gravity O having four faces 9. The optical instrument 2 is mounted on one of the faces 9 of the body 3 so as to to be able to direct its V-axis of aim towards the Earth T for a shot.
    Two repositories are defined for satellite 1.
    A first reference OXYZ is related to the orbit A of the satellite 1 at a given point, here the center O of gravity of the body 3, and is called local orbital reference. The OXYZ reference local orbital is orthogonal, and comprises three axes: an axis parallel to the speed vector of the satellite 1 on its orbit A noted X and called speed axis, an axis perpendicular to the plane of the orbit A noted Y and called orbital axis, and an axis pointing to the main focus of the orbit A noted Z, that is to say towards the Earth T, and called axis Nadir.
    In what follows, the attitude of the satellite 1 is defined as the movement of the satellite 1 in the local orbital reference OXYZ.
    A second Oxyz repository is linked to the body 3 of the satellite 1 and its faces, and is called satellite reference. The Oxyz satellite repository is orthogonal. It comprises a first axis x, a second axis y and a third axis z. The third axis z is parallel to the axis (V) of sight of the optical instrument (2) and is oriented in the same direction.
    In the example of Figure 1, the body 3 is parallelepiped, each axis is perpendicular to a face of the body 3 of the satellite.
    In FIG. 1, the second Oxyz repository has been represented in a position in which it is merged with the first OXYZ repository. The optical instrument 2 is fixed relative to the body of the satellite 1, that is to say that its axis V of sight does not move relative to the Oxyz satellite reference. For purposes of simplifying the description, the reference Oxyz satellite is in the following such that the z axis is coincident with the V axis of sight of the optical instrument 2, and the x axis coincides with the X axis speed of the local orbital repository.
    The satellite 1 also comprises a solar generator 4. The solar generator 4 is also fixed relative to the body 3 of the satellite, that is to say that it is fixed with respect to the satellite reference system. More precisely, the solar generator 4 has at least one functional surface, that is to say a surface equipped to receive the solar energy and transform it into energy that can be used by the satellite 1, oriented along a normal N. For example, the solar generator 4 is a solar panel. In what follows, we will speak of sunshine to define the power received by the solar generator 4 and sun rays.
    According to the example of FIG. 1, the normal N of the solar generator 4 is parallel to the Z axis and oriented in the opposite direction to that of the optical V axis of the optical instrument 2. The normal N at the surface of the solar generator comprises at least one component perpendicular to the axis X speed.
    The satellite 1 further comprises an attitude control device 100 and a control unit 102 connected to this attitude control device 100.
    The attitude control device comprises, in particular, inertial actuators. These inertial actuators make it possible to set the body 3 of the satellite 1 into motion in the local orbital reference frame Oxyz. These are for example reaction wheels or CMGs.
    The inertial actuators of the attitude control device 100 are arranged perpendicular to each other on the main body 3 of the satellite. Some inertial actuators make it possible to rotate the satellite 1 around the first axis x. They are called actuators along the first axis x. Some inertial actuators make it possible to rotate the satellite 1 around the second axis y. They are called actuators along the second axis y. And finally other inertial actuators make it possible to rotate the satellite 1 around the third axis z. They are called actuators along the third axis z.
    According to the invention, the inertial actuators along the second axis y and / or the actuators along the third axis z have a maximum torque capacity of less than 40% of the maximum torque capacity of the actuators along the first axis x. Preferably, the inertial actuators along the second axis y and / or along the third axis z have maximum torque capacities less than 30% of the maximum torque capacity of the actuators along the first axis x.
    In the present patent application, torque capacity is the maximum torque that an inertial actuator can generate to rotate the satellite.
    Advantageously, the inertial actuators along the second axis y and / or the actuators along the third axis z mounted in the satellite according to the invention are therefore smaller and lighter than the inertial actuators according to these axes mounted in the satellites of the state of the art.
    More generally, the three actuators can be arranged differently, without being allocated individually to a rotation along the x, y or z axes. Thus, one or more of the three actuators may not be arranged on the x, y and z axes.
    A larger number of actuators can be used for redundancy purposes. The maximum torque capacity along the second axis y and / or along the third axis z, however, will remain 40% less than the capacity along the first axis x, which allows the use of a cluster of actuators lighter than the state of the art.
    Preferably, the maximum torque capacity along the second axis y and / or along the third axis z is 30% lower than the capacity along the first axis x. The control unit 102 includes a memory and a computing unit. This is for example a processor. The control unit 102 is able to implement the guiding method according to the invention.
    With reference to FIG. 2A, the guiding method comprises a step 104 for transmitting guiding commands from the control unit 102 to the attitude control unit 100.
    These guidance commands are defined by the control unit from a maneuvering plane transmitted to the satellite by an operator located in a station on the Earth or by control laws and / or tables defined according to the position from the sun relative to the orbit A. These control laws or these tables are prerecorded in the memory of the control unit or transmitted to it by a ground operator.
    According to the method of the invention, the guidance commands comprise only commands for rotating the satellite 1 around the X-axis speed, that is to say roll commands.
    The rotation commands of the attitude control command are zero around the orbit axis Y and the Nadir Z axis. In practice, it is possible that the angle of rotation is non-zero around the orbit axis Y or Nadir Z axis due to residual errors of the attitude control law and the presence of couples disruptors. Nevertheless, in such a case, these angles are negligible. These angles of rotation are for example less than 2 degrees and preferably less than 1 degree. The angle of rotation around the orbit Y axis and the Nadir Z axis in the orbital reference system is thus maintained substantially zero. The transmission step 104 is continued by a step 106, during which the satellite pivots only about the speed X axis to point the viewing axis V towards a region to be imaged. In this position, the optical instrument images a region 7, by one or more shots.
    The surface of the region to be imaged 7 is limited by the field of view of the optical instrument 2.
    In a step 108, the control unit 102 transmits a rotation command only about the speed axis X to image another portion of the earth.
    During a step 109, the satellite 1 is pivoted only around the axis X speed, so that the line of sight V is directed towards the other portion of the earth, so that the optical instrument image a other region 7.
    During a step 110, the control unit 102 again transmits a guidance command to the attitude control device 100.
    This guidance command is able to minimize the angle between the normal N of the functional surface of the solar generator 4 and the direction R of the solar rays to obtain maximum sunlight. This guidance command has only one rotation around the speed X axis.
    During a step 112, the satellite 1 pivots only about the speed axis X. The satellite 1 is then in the waiting status.
    FIG. 2B schematically represents the attitude of the satellite 1 in the particular case of an initial state in which the Oxyz satellite reference is merged with the local orbital reference frame OXYZ. In this diagram, there is shown a second position of the satellite 1 in which it has rotated by an angle δ about the X axis speed from the initial position, so that the y and z axes respectively of the satellite reference frame are displaced by an angle δ with respect to the Y and Z axes respectively of the local orbital reference frame. The x-axis of the satellite reference is always coincident with the X axis of the local orbital frame, since no rotation around the Y and Z axes has taken place.
    Thus, according to the method according to the invention, only rotations around the axis X speed are performed by the satellite.
    The result is a satellite of a simplified design. In particular, the attitude control device only needs to be operational in rotation about the X axis speed for the guidance of the attitude of the satellite.
    According to one embodiment of the invention, the satellite 1 is part of a satellite constellation, that is to say a group of satellites 1 intended to work together in the context of a mission. The satellites of the constellation whose attitude is guided by the method according to the invention may be identical, facilitating mass production without increasing costs.
    For example, in FIG. 3, three satellites 1 of the same constellation have been represented, each moving in an orbit A, B and C respectively. The attitude of these satellites is guided by the control method according to FIG. invention. By pivoting the three satellites 1 only around their axis X speed over defined angular ranges, and considering their displacement on their orbit A, B and C, the line of sight V of the optical instrument of each satellite is able to scan a region of the ground 5, so that together, the three satellites 1 can potentially take a view of the ground of a region 6 more important.
    FIG. 4 schematically shows various regions 7 of the ground that a satellite 1 moving in a orbit A corresponding to the ground at a trajectory Tr can take an image. By pivoting about the axis X speed, the line of sight of the optical instrument 2 follows a trajectory tr, included in a ground strip of a width L. When two regions 7, 7 'of the ground to be taken in image are side by side, on both sides of the trajectory Tr on the ground of a satellite 1, the satellite 1 takes a first region 7 of the ground and another satellite 1 of the same constellation can be charged with the shooting of the other region 7 '. In the example of Figure 3 the satellites are on the same orbit, but they can be on different orbits.
    It is also possible to take the same region or area from different angles, for example to obtain a stereo image.
    Thus, by moving the satellite 1 into the operating status only around the X-axis speed, and considering a constellation of satellites, the results of the shots are at least equivalent to those of the state of the art.
    Since the optical instrument 2 does not need to include moving parts, the quality of the shots is increased. Indeed, the line of sight and the field of view of the optical instrument 2 are not displaced relative to the body 3 of the satellite. This stability improves the quality and therefore the precision of the shots.
    When the satellite 1 is in the waiting state, the body 3 of the satellite 1 is rotated about the speed axis X to orient the solar generator 4 towards the R-rays of the Sun. For example, the best position around the X axis is calculated so that the angle between the normal N of the surface of the solar generator 4 and the direction R of the sun's rays is as small as possible. It is therefore not necessary to provide in advance a position of the solar generator 4 on the body 3 of the satellite 1 according to its orbit, since the attitude guide method optimizes the sunshine of the generator 4 solar by rotation around the X axis speed. Thus the same design can be maintained for a satellite constellation in sun-synchronous orbits (SSO) with different local hours at the upstream node.
    According to an example (FIGS. 6 and 7), the solar generator 4 comprises at least one solar panel 8 whose functional surface, that is to say the surface covered with solar cells, is oriented by the normal N. Thus, from of the position of Figure 1 for example, the satellite 1 is rotated about the axis X speed to orient the normal N almost parallel to the radius of direction R. Depending on the position of the satellite on the orbit A, the orientation the direction R of the sun's rays varies. However, the solar generator 4 is fixed relative to the body 3 of the satellite, so that the normal N of the solar generator is almost never perfectly parallel to the direction R of the sun's rays. The solar generator 4 may then comprise a plurality of solar panels, each solar panel comprising a functional surface whose normal comprises at least one component perpendicular to the axis X speed. The normal N of the functional surface of the solar generator 4, taken into account for the calculation of the optimization in the waiting status, can then be an average normal to the normals of the solar panels, or be confused with at least one of the normal solar panels.
    For example, the solar generator 4 may comprise two panels, respectively 8a and 8b, each having a functional surface with a normal, respectively Na and Nb. The two normal Na and Nb are preferably parallel to each other. For example, the two panels 8a, 8b are hung on either side of a satellite face. At launch they will preferably be in folded configuration and deployed in orbit. The two panels can be aligned with this face of the satellite, to form a plane (Figure 8), or be inclined relative to this face (Figures 9 and 10). In general, each face of the solar panel 8 is always oriented so as to never have its normal oriented in a direction opposite to that of the normal of another panel, so that it is possible, by rotation around the X axis speed, to find a position in which none of the panels 8a, 8b is in the shadow of the Sun's rays.
    Thanks to the positioning in the standby status of the solar generator 4 so as to optimize its sunshine, the power received by the solar generator 4 is on average greater than that received by a solar generator that would have been mounted on the body of a satellite with optimized initial calibration. For example, it has been established that for a solar generator 4 having a functional surface of 2.4 m2 mounted on a satellite 1 of sun-synchronous orbit, whose attitude is guided by the method according to the invention, that is that is to say only in rotation around the axis X speed, the available power at the level of the satellite is between 230W and 350W, according to the local time (LTAN) considered and the day in the year, whereas for a solar generator fixed on the body of the satellite so as to optimize its sunshine when pointing the optical axis of the instrument to the Earth, the power generated is between 205W and 310W. Thus, thanks to the method according to the invention, the maximum power generated is increased between 12% and 18%.
    Thus, the method for controlling the attitude of a satellite according to the invention allows for any satellite, independently of both its orbit characterized by its local time LTAN and the day in the year (solstices and equinoxes) in the case of a sun-synchronous orbit, to adapt the position of the satellite in the waiting status to maximize the sunshine of its solar generator. Specifically, the LTAN local time characterizes the inclination of the orbit of the satellite around the Earth T. Thus, for satellites whose orbit is characterized by a different LTAN local time, the solar generator 4 must be oriented differently to optimize its sunshine in the waiting status. With guidance of the attitude by rotation only around the X-axis speed, it is sufficient to adapt the orientation of the satellite around the X-axis speed to maximize solar group 4 sunshine. The attitude guidance method according to the invention makes it possible to have a satellite with the expected functionalities with increased performances, for a simpler design. Since the optimization of the solar generator 4 sunshine is regulated solely by rotation around the X axis speed, several satellites 1 having different orbits can have the same design, and in particular solar panels 8 mounted identically on the satellite body.
    For a constellation in sun-synchronous orbits, it may be advantageous to adapt the angle of the solar generator according to the local time so as to reduce the roll attitude deflections.
    FIGS. 11 and 12 are graphs each illustrating the attitude of a satellite in the waiting status, guided by the method according to the invention, the abscissa axis bearing the true anomaly of the satellite orbiting the Earth, the y-axis carrying the rotation angle of the satellite around the X-axis speed, the rotation around the Y and Z axes being almost zero.
    On the graph of FIG. 11, it is considered a satellite 1 whose solar generator 4 has its normal N opposite the V axis of aiming of the instrument 2 - it is then called a solar generator at the zenith - as illustrated in FIG. 1 in particular, circulating in a sun-synchronous orbit characterized by its local time (LTAN). From an initial position (true anomaly equal to 0), the satellite is rotated about its axis X speed by an angle of 22.5 ° corresponding to the inclination of the satellite orbit for the local time at 10:30. As the satellite moves on its orbit and the true anomaly increases, the angle of rotation around the X-axis speed increases, to optimize the sunshine of the solar generator 4. However, preferably, once the rotation about the X axis has reached a determined value, in this case 50 ° in the example of Figure 11, the rotation can be restrained, depending on the allowable stresses at the satellite. The actual curve of the attitude of the satellite as a function of the true anomaly, in full line, then deviates from the theoretical curve, in discontinuous lines, the latter not taking into account the clamping. This clamping makes it possible in particular to save other parameters of the satellite, and to avoid the glare of sensitive equipment, such as stellar sensors. When the satellite 1 is in the shadow of the Earth (from 110 ° true anomaly in Figure 11), its illumination by the Sun is zero, so that to seek to illuminate the solar generator 4 is no longer adapted, the recovered power being zero. The satellite 1 can then be rotated about the axis X speed so as to reduce the angle of rotation around the axis X speed 0 °, corresponding to a position in which the V axis of sight points in the direction of center of the Earth. Here again, the real curve of the attitude of the satellite as a function of its true anomaly is removed from the theoretical curve. This pointing allows for example to maintain a thermal stability of the optical instrument 2 and recover some of the heat of the Earth. Then, once the satellite passes again in an illuminated area (at about 245 ° true anomaly in Figure 11) by the Sun, the rotation around the X axis speed continues, first with a clamping at 50 ° for the same reasons as those described above, then the real curve joins the theoretical curve to optimize the sunshine of the solar generator 4.
    On the graph of FIG. 12, this time it is considered another example of a satellite 1 whose solar generator 4 has its normal N forming an angle of 45 ° with respect to the normal of the preceding example - then we speak solar generator set at 45 ° - in particular, circulating in an orbit characterized by its local time (LTAN) at 9 o'clock. In the initial position (true anomaly equal to 0 °), the rotation angle around the X-axis speed is zero, because the 45 ° setting is optimized in this case for the 9h LTAN orbit. The evolution of the attitude of the satellite around the axis X speed is as before, with a clamping at 50 ° and a setting to 0 ° of the angle of rotation around the axis X speed when the satellite is in the shadow of the Earth to preserve the optical instrument.
    Guiding the attitude of the satellite only in rotation about the speed X axis thus simplifies the design of the satellite 1.
    Thus, for example, the body 3 of the satellite always has the same front face 9, that is to say the face perpendicular to the axis X speed, so that it is possible to reduce the surface of the front face to minimize drag, without affecting the rest of the body 3 of the satellite 1.
    The simplicity of the design of the satellite whose equipment does not include a mobile part and which is provided with a maneuvering device according to the simplified Y and Z orbit axes, has the effect that the satellite can be considerably lightened. .
    Thus, the inertia of the satellite 1 rotating around the X-axis speed is reduced, allowing it to go from the shooting status to the waiting status quickly with actuators of lesser capacity than for satellites maneuvering the vehicle. state of the art.
    According to an exemplary embodiment of the satellite as shown schematically in FIG. 13, the V-axis of sight of the optical instrument 2 is oriented towards a launcher interface device 10, such as a ring, intended to cooperate with a device complementary interface of a launcher and / or another satellite. The launcher interface device 10 is generally a ring. Schematically, the body of the satellite is represented as being reduced to a plate 3 connected to the ring 10 by a structure 11. The optical instrument 2 is fixed rigidly on the plate 3 inside the structure 11. The axis The aiming V of the optical instrument 2 is then oriented from the upper end to the lower end of the intermediate structure 11 towards the starter interface ring. The intermediate structure 11 can form an enclosure within which the optical instrument 2 is housed. The satellite 1 can be mounted in a launcher by putting the ring 10 in cooperation with a complementary ring of a launcher, or it can be stacked on a satellite by putting the ring 10 in cooperation with a complementary device of the other satellite.
    This design makes it possible to have a more compact satellite 1, further reducing its inertia for guiding its attitude. This design also serves to attenuate the vibrations transmitted from the launcher by the satellite interface ring 10 to protect the optical instrument 2. Indeed, the optical instrument 2 is remote from the interface ring 10 by the intermediate structure 11, the latter absorbing the vibrations at least in part before they reach the body 3 and the optical instrument 2.
    In addition, this targeted design makes the satellite particularly suitable for stacking satellites in a launcher for multiple launch, for example as part of a constellation of satellites. However, the place under a cap in a launcher is generally reduced, and limits the number of satellites that can be stacked. Due to the compactness, for a given launcher, a stack of satellites according to this design can comprise a larger number of satellites than for satellites of the state of the art.
    权利要求:
    Claims (13)
    [1" id="c-fr-0001]
    1, - Method for controlling the attitude attitude of a satellite (1) with respect to an orthogonal orbital reference (OXYZ) comprising a speed axis (X), an orbital axis (Y) and an axis (Z) ) says Nadir, along a portion of orbit (A) around the Earth (T), said orbit portion (A) being illuminated by solar radiation; the satellite moving in the direction of the speed axis (X), the satellite (1) comprising a main body (3), an optical instrument (2) having a fixed aiming axis (V) relative to the body (3) ) main, at least one solar generator (4) fixed relative to the main body (3) defining a functional surface whose normal (N, Na, Nb) has at least one component perpendicular to the axis (X) speed, to at least one attitude control device (100) and a control unit (102) connected to the attitude control device (100), said method comprising a guide control transmission step (104) of the control unit (100). controlling (102) the attitude control unit (100) to point the optical instrument aiming axis (V) towards regions (7) to be imaged or to direct the normal (N, Na, Nb) to the functional surface in the direction of solar radiation, characterized in that the guiding commands are a However, only the satellite rotation commands around the speed axis (X), the angle of rotation about the orbit axis (Y) and the Nadir axis (Z) in the orbital frame being kept substantially impaired.
    [2" id="c-fr-0002]
    2, - The method of claim 1, wherein the guidance commands comprise commands adapted to rotate the satellite (1) about the axis (X) speed over angular ranges to scan with the axis (V) of aimed at a portion of the earth.
    [3" id="c-fr-0003]
    3, - Method according to claim 1 or claim 2, implemented by several satellites (1).
    [4" id="c-fr-0004]
    4, - Method according to any one of the preceding claims, wherein the normal (N, Na, Nb) at the functional surface of the solar generator (4) is parallel to the axis (V) of sight of the instrument ( 2) optical and oriented in the opposite direction.
    [5" id="c-fr-0005]
    5, - A method according to any one of the preceding claims, wherein, the angles of rotation about the axis (X) speed are limited to a predetermined clamping angle, said guide angle being defined with respect to the axis (Y) orbital.
    [6" id="c-fr-0006]
    6, - The method of claim 5, wherein the clamping angle is 50 °.
    [7" id="c-fr-0007]
    7, - Method according to any one of the preceding claims, wherein when the illumination of the satellite (1) is zero, the satellite (1) is rotated about the axis (X) speed so as to point the axis (V) aiming towards Earth (T).
    [8" id="c-fr-0008]
    8, - Method according to any one of claims 1 to 7, wherein the line of sight (V) is perpendicular to the speed axis (X).
    [9" id="c-fr-0009]
    9, - Satellite (1) comprising a main body (3), an optical instrument (2) whose aiming axis (V) is fixed relative to the main body (3), at least one fixed solar generator (4) with respect to the main body (3), at least one attitude control device (100) and a control unit (102) connected to the attitude control device (100), the control unit (102) being adapted to perform the guidance control method according to any one of the preceding claims, the attitude control device (100) is adapted to rotate the satellite (1) about a first axis (x), a second axis (y) and a third axis (z), said first (x), second axis (y) and third axis (z) being perpendicular to each other, said third axis (z) being parallel to the axis (V) of sight of the instrument (2) optical and oriented in the same direction, characterized in that the torque capacity of the con attitude control (100) along the second axis (y) and / or along the third axis (z) is less than 40% of the torque capacity along the first axis (x).
    [10" id="c-fr-0010]
    10, - Satellite (1) according to the preceding claim, comprising an interface device (10) intended to cooperate with a complementary interface device of a launcher or a satellite, and comprising an intermediate structure (11) connecting the body (3) of the satellite (1) to the interface device (10), the aiming axis (V) of the optical instrument (2) being directed towards the interface device (10).
    [11" id="c-fr-0011]
    11, - Satellite (1) according to any one of claims 9 and 10, wherein the normal (N, Na, Nb) to the functional surface of the solar generator (4) is parallel to the axis (V) of sight the instrument (2) is optical and oriented in the opposite direction.
    [12" id="c-fr-0012]
    12, - plurality of satellites shaped according to any one of claims 9 to 11 and intended to operate in a constellation, said plurality of satellites being adapted to be guided in orbit by the method according to any one of claims 1 to 8.
    [13" id="c-fr-0013]
    13, - Computer program product characterized in that it comprises a set of program code instructions which when executed by a processor, implement a method according to claims 1 to 8.
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    同族专利:
    公开号 | 公开日
    EP3390230B1|2019-06-12|
    US20190033891A1|2019-01-31|
    FR3047813B1|2019-11-08|
    EP3390230A1|2018-10-24|
    WO2017140972A1|2017-08-24|
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    WO2009150081A1|2008-06-09|2009-12-17|Astrium Sas|Method for controlling satellite attitude, and attitude-controlled satellite|
    EP2489593A1|2011-02-21|2012-08-22|European Space Agency|Earth observation satellite, satellite system, and launching system for launching satellites|
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    US10832054B2|2018-10-15|2020-11-10|National Applied Research Laboratories|Guidance seeking device for a satellite and method for enhancing the performance of the same|
    CN109901581A|2019-03-15|2019-06-18|智久机器人科技有限公司上海分公司|A kind of scaling method and spin motion control method of AGV vehicle spin angle|
    CN111366986A|2020-03-24|2020-07-03|中国科学院微小卫星创新研究院|Space debris observation system and method|
    CN111619825B|2020-04-29|2021-12-21|北京航空航天大学|Cross-cut formation method and device based on star-sail rope system|
    WO2022010819A1|2020-07-10|2022-01-13|Star Mesh LLC|Data transmission systems and methods for low and very low earth orbit satellite communications|
    CN111959830B|2020-08-24|2021-10-15|中国科学院微小卫星创新研究院|Thermal control system and method for satellite high-precision optical load mounting platform|
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    2017-08-18| PLSC| Publication of the preliminary search report|Effective date: 20170818 |
    2018-02-26| PLFP| Fee payment|Year of fee payment: 3 |
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    2021-11-12| ST| Notification of lapse|Effective date: 20211005 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1651237A|FR3047813B1|2016-02-16|2016-02-16|METHOD FOR CONTROLLING SATELLITE ATTITUDE GUIDANCE, SATELLITE, SATELLITE PLURALITIES AND ASSOCIATED COMPUTER PROGRAM|
    FR1651237|2016-02-16|FR1651237A| FR3047813B1|2016-02-16|2016-02-16|METHOD FOR CONTROLLING SATELLITE ATTITUDE GUIDANCE, SATELLITE, SATELLITE PLURALITIES AND ASSOCIATED COMPUTER PROGRAM|
    US16/072,996| US20190033891A1|2016-02-16|2017-02-13|Method for controlling the attitude guidance of a satellite, satellite, pluralities of satellites, and associated computer program|
    EP17708861.4A| EP3390230B1|2016-02-16|2017-02-13|Method for controlling the guidance of attitude of a satellite, satellite, pluralities of satellites, and associated computer programme|
    PCT/FR2017/050318| WO2017140972A1|2016-02-16|2017-02-13|Method for controlling the guidance of attitude of a satellite, satellite, pluralities of satellites, and associated computer programme|
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