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    • 专利摘要:
    The present invention relates to a method (50) for controlling a phase of acquisition of the Sun by a spacecraft with a non-zero kinetic momentum of axis DH, said spacecraft comprising a solar generator rotatable about an axis Y, comprising steps of: - (51) controlling the spacecraft actuators so as to place said spacecraft in an intermediate orientation in which the Y axis is substantially orthogonal to the axis DH, - (52) control of the solar generator so as to orient said solar generator towards the Sun, - (53) control of actuators of the spacecraft so as to reduce the kinetic moment of said spacecraft, - (54) control actuators of the spacecraft so as to place said spacecraft in an acquisition orientation in which the Y axis is substantially orthogonal to the DS direction of the Sun relative to the spacecraft.
    公开号:FR3013685A1
    申请号:FR1361581
    申请日:2013-11-25
    公开日:2015-05-29
    发明作者:Nicolas Cuilleron；Jean Sperandei；Philippe Laurens
    申请人:Astrium SAS；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD The present invention belongs to the field of spacecraft control, and more particularly relates to a method and a device for controlling a phase of acquisition of the Sun by a spacecraft 5 comprising at least one mobile solar generator around the planet. a Y axis. By "acquisition phase of the Sun" is meant a phase during which the spacecraft is controlled so as to place said spacecraft in an acquisition direction in which the Y axis is substantially orthogonal to the direction Ds of the Sun with respect to the spacecraft. In this acquisition orientation, the solar generator can be directed to the Sun, whose rays then have a normal incidence on said solar generator. STATE OF THE ART A phase of acquisition of the Sun is for example carried out by a spacecraft, such as a satellite, after separation with a launcher of said satellite. Such a phase of acquisition of the Sun is designated in the following description by "initial acquisition phase of the Sun". Indeed, after separation with the launcher, the satellite is autonomous and must ensure its electric autonomy by seeking to point its solar generators to the Sun to provide electrical power to a platform 20 of said satellite, and to recharge batteries of said satellite. Conventionally, the batteries are charged before launching the satellite and, after the separation and before the solar generators are deployed and directed to the Sun, the electrical power necessary for the operation of the satellite platform is provided by said batteries. This initial acquisition phase of the Sun is made difficult because at the moment of separation with the launcher, the satellite can be released with a high rotational speed: up to 3 ° / s typically in LEO orbit (" Low Earth Orbit "), and of the order of 1 ° / s in orbit GEO (" Geostationary Orbit "). The initial acquisition phase of the Sun therefore comprises a step of reducing the speed of rotation of the satellite. This reduction in the speed of rotation of the satellite must be short so as not to risk completely discharging the batteries of said satellite. Thus, the satellite comprises actuators used to reduce the kinetic moment of the satellite after separation with the launcher. It is particularly known, for satellites in LEO orbit, to use magneto-couplers exploiting the Earth's magnetic field to form couples adapted to reduce the kinetic moment of the satellite. For satellites in MEO ("Medium Earth Orbit") or GEO (or GTO - "Geostationary Transfer Orbit"), for which the Earth's magnetic field is negligible, chemical thrusters are generally used. Such chemical thrusters are advantageous in that they can be implemented immediately after separation with the launcher, in that they do not require high electrical power, and in that they can apply significant torques on the satellite which makes it possible to reduce the kinetic momentum and to place the satellite in the acquisition orientation in a very short time (of the order of a few minutes). After the kinetic moment of the satellite has been sufficiently reduced and the satellite has been placed in the acquisition orientation, the solar generators are deployed and directed towards the Sun to power the platform of said satellite and recharge the batteries. It is currently envisaged that future satellites will no longer be equipped with chemical thrusters, but only with electric (plasma) thrusters. The initial acquisition phase of the Sun would then become problematic, especially for satellites in the MEO / GEO / GTO orbit, since electric thrusters can not replace chemical thrusters to reduce the kinetic momentum of the satellite after separation. Indeed, electric thrusters can not be used during the first days after separation with the launcher. In addition, the thrusters require a high electrical power (of the order of a few kW) and can only apply low torques on the satellite (of the order of 0.2 N.m). Consequently, the reduction of the kinetic moment of the satellite by means of electric thrusters would require a significant electric power during a long duration, which is incompatible with the batteries embarked on the current satellites, which have a capacity of the order of 10 kW. h. Such a capacity allows to power the electric thrusters only about an hour, which is insufficient to sufficiently reduce the kinetic moment of the satellite, given the couples that can be formed by the current electric thrusters. Current satellites are generally equipped with electric inertial actuators, such as flywheels (reaction wheels, gyroscopic actuators), which are implemented to control the attitude of said satellite along three axes. Such inertial actuators have a generally insufficient capacity to compensate for the kinetic moment of the satellite after separation, but can be implemented after the other actuators (magnetocouplers, chemical thrusters, etc.) have sufficiently decreased said kinetic moment of said satellite . An alternative could be to size the electric inertial actuators so that they have a capacity adapted to compensate only the kinetic moment of the satellite after separation. However, this would cause a significant increase in the mass and volume of said inertial actuators. For example, the initial kinetic moment of a satellite after launch can be of the order of 500 Nms to 1000 Nms, while the capacity of the inertial actuators embedded in a satellite is generally of the order of 50 Nm. s to 100 Nm.s. In addition, even if the inertial actuators are sized to have a capacity of 1000 Nm.s, particularly unfavorable separation conditions could still transfer to the satellite an initial kinetic moment greater than 1000 N.m.s, outside the capacity of the inertial actuators. It should be noted that an acquisition phase of the Sun can also be performed when the satellite is in survival mode. The previous observations remain valid for such a phase of acquisition of the Sun in survival mode. DISCLOSURE OF THE INVENTION The present invention aims to remedy all or part of the limitations of the solutions of the prior art, including those described above, by proposing a phase of acquisition of the Sun that can be implemented y 30 understood by spacecraft without chemical propellants. For this purpose, and according to a first aspect, the invention relates to a method of controlling a phase of acquisition of the Sun by a spacecraft with a non-zero kinetic momentum of axis DH, said spacecraft comprising a body, a solar generator and a drive mechanism of said solar generator rotating about a Y axis. According to the invention, the control method comprises steps of: - controlling the actuators of the spacecraft so as to place said machine space in an intermediate orientation, with respect to the kinetic moment, in which the Y axis is substantially orthogonal to the axis KD of the kinetic moment, - control of the drive mechanism of the solar generator, previously at least partially deployed, so as to directing said solar generator towards the Sun, - controlling the actuators of the spacecraft so as to reduce the kinetic moment of said spacecraft, - controlling the actuators of the spacecraft so as to place said spacecraft in an acquisition orientation, with respect to the Sun, in which the Y axis is substantially orthogonal to the direction Ds of the Sun with respect to the spacecraft. The intermediate orientation of the spacecraft makes it possible to ensure that there is always an orientation of the solar generator to ensure the electrical autonomy of said spacecraft. By at least partially deploying the solar generator before reducing all or part of the kinetic moment of the spacecraft, the constraints related to the electric autonomy of the spacecraft are released. Said reduction in kinetic moment can in particular be spread over time and / or implement electric actuators. In particular embodiments, the Sun acquisition phase control method may further include one or more of the following features, taken alone or in any technically possible combination. In a particular mode of implementation, the placement of the spacecraft in the intermediate orientation uses inertial actuators 30 of said spacecraft, preferably gyroscopic actuators and / or reaction wheels. In a particular mode of implementation, the placement of the spacecraft in the acquisition orientation uses inertial actuators of said spacecraft, preferably gyroscopic actuators and / or reaction wheels. In a particular mode of implementation, the inertial actuators are implemented, to place the spacecraft in the acquisition orientation, when the kinetic moment of the spacecraft becomes less than a predefined threshold value. In a preferred mode of implementation, the reduction of the kinetic moment of the spacecraft, previously placed in the intermediate orientation, uses electric thrusters of said spacecraft.
    [0002] In a particular mode of implementation, only electric actuators are implemented during the acquisition phase of the Sun. In a particular mode of implementation, the maximum axis of inertia of the spacecraft, when the solar generator of said spacecraft is deployed, being orthogonal to the Y axis, the solar generator is at least partially deployed beforehand. placement of the spacecraft in the intermediate orientation and, during the placement of said spacecraft in said intermediate orientation, a nutation damping is performed. In a particular mode of implementation, the minimum axis of inertia of the spacecraft being orthogonal to the axis Y, it is carried out, during the placement of the spacecraft in the intermediate orientation, a control of nutation aiming at aligning the minimum axis of inertia of said satellite with the axis DH of the kinetic moment. In a particular mode of implementation, the orientation of the solar generator during the reduction of the kinetic moment is determined as a function of the angle between the direction Ds of the Sun and the axis DH of the moment kinetic, called "angle of misalignment 0 ". In a particular mode of implementation, the orientation of the solar generator is kept fixed relative to the body of the spacecraft during the reduction of the kinetic moment.
    [0003] In a particular mode of implementation, the solar generator is maintained, during the reduction of the kinetic momentum, so that a photosensitive surface of said solar generator is: - substantially parallel to the axis DH of the moment kinetic when 'cos ei <Vs1, Vs1 being a predefined threshold value, substantially orthogonal to the axis DH when 'cos ei> Vs1, said photosensitive surface being oriented towards the Sun. In a particular mode of implementation, the spacecraft comprising two solar generators mounted on respective drive mechanisms adapted to rotate said solar generators around respective parallel Y axes, said solar generators are maintained, during the reducing the kinetic momentum, so that respective photosensitive surfaces of said solar generators are substantially parallel to the axis DH of the kinetic moment and with respective pointing directions opposite when 'cos ei <Vs2, Vs2 being a predefined threshold value. In a particular mode of implementation, the phase of acquisition of the Sun by the spacecraft is controlled remotely by a control device, control signals being successively determined and sent to the spacecraft by said control device . According to a second aspect, the invention relates to a computer program product comprising a set of program code instructions which, when executed by a processor, implement a control method of an acquisition phase. of the Sun by a spacecraft according to any one of the embodiments of the invention. According to a third aspect, the invention relates to a device for controlling a phase of acquisition of the Sun by a spacecraft with a non-zero kinetic momentum of axis DH, said spacecraft comprising a body, actuators, a solar generator and a mechanism for driving said solar generator in rotation about a Y axis, said control device comprising means configured to control the acquisition phase of the Sun according to any of the embodiments of the invention . According to a fourth aspect, the invention relates to a space system comprising a spacecraft comprising a body, actuators, a solar generator and a drive mechanism of said solar generator rotating about a Y axis, said system further comprising a control device for a phase of acquisition of the Sun by said spacecraft according to any one of the embodiments of the invention.
    [0004] In a preferred embodiment, the spacecraft is devoid of a chemical propellant. PRESENTATION OF THE FIGURES The invention will be better understood on reading the following description, given by way of non-limiting example, and with reference to the figures which show: FIG. 1: a schematic representation of a space system comprising a spacecraft having to execute a phase of acquisition of the Sun, - Figure 2: a diagram illustrating the main steps of a phase acquisition method of acquisition of the Sun, - Figure 3: a curve representing the average insolation of solar generators a spacecraft placed in an intermediate orientation during a phase of acquisition of the Sun, - Figure 4: a schematic representation of an example of implementation of a phase of acquisition of the Sun. In these figures, identical references from one figure to another designate identical or similar elements. For the sake of clarity, the elements shown are not to scale unless otherwise stated.
    [0005] DETAILED DESCRIPTION OF EMBODIMENTS The present invention relates to the phases of acquisition of the Sun by a spacecraft comprising at least one mobile solar generator around a Y axis of the spacecraft. By "phase of acquisition of the Sun" is meant a phase during which the spacecraft is controlled so as to place said spacecraft in an acquisition orientation in which the Y axis is substantially orthogonal to the direction Ds of the Sun in relation to the spacecraft. In particular, the present invention finds a particularly advantageous application in the case of a phase of initial acquisition of the Sun by the spacecraft (immediately after separation with a launcher of the spacecraft) or a phase of acquisition of the Sun in the survival mode of the spacecraft. In the remainder of the description, one places oneself in a nonlimiting manner in the case where the spacecraft is a satellite 10 and in the case of a phase of initial acquisition of the Sun by the satellite 10. It should be noted in in addition that the invention is applicable regardless of the LEO orbit, MEO, GEO, etc., on which the satellite 10 is intended to carry out its mission. The present invention finds a particularly advantageous application, particularly in the case of a satellite 10 placed in a GTO orbit by a launcher and intended to perform its mission in GEO orbit. FIG. 1 schematically represents a space system comprising a satellite 10 to be executed. an initial acquisition phase of the Sun 60, after separation with a launcher (not shown in the figures).
    [0006] As illustrated by FIG. 1, the satellite 10 comprises a body 11, and two solar generators 12 on either side of the body 11, and a set of actuators (not shown in the figures) adapted to control the orbit of the satellite and to control the attitude of the satellite, preferably along three axes. The solar generators 12 comprise respective photosensitive surfaces, for example flat surfaces, which when directed towards the Sun generate electrical power. In order to control the orientation of said photosensitive surfaces with respect to the Sun, the solar generators 12 are rotatable about a same axis Y, around which said solar generators 12 are driven by respective drive mechanisms.
    [0007] It should be noted that the solar generators 12 are shown deployed in FIG. 1 but that, immediately after separation with the launcher, said solar generators 12 are generally folded. In addition, the satellite 10 illustrated in FIG. 1 comprises two solar generators 12, but the invention is applicable as soon as said satellite 10 comprises at least one solar generator 12 mobile in rotation about a Y axis. The space system comprises also a device 20 for controlling the initial acquisition phase of the sun 60 by the satellite 10. In the nonlimiting example illustrated in FIG. 1, the control device 20 is remote from the satellite 10, and corresponds to a ground station on the surface of the Earth 61.
    [0008] More generally, the control device 20 can be embedded in the satellite 10 or in one or more devices remote from said satellite 10. Moreover, nothing further excludes, according to other examples, having a distributed control device 20. between the satellite 10 and one or more other equipment remote from said satellite 10. In the example illustrated in FIG. 1, the control device 20 thus remotely controls the initial acquisition phase of the Sun by the satellite 10, and the device 20 and the satellite 10 comprise for this purpose the respective conventional means of remote communication. The control device 20 is adapted to determine control signals of the initial acquisition phase of the Sun, which are sent to the satellite 10. Said control signals are for example determined as a function of measurement signals received from the satellite 10 which are determined by different sensors (gyroscope, gyrometer, accelerometer, star sensor, etc.) of said satellite 10, the control device 20 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 various steps of a control method. In a variant, the control device 20 also comprises one or more programmable logic circuits, of the FPGA, PLD, etc. type, and / or specialized integrated circuits (ASIC) adapted to implement all or part of said process steps 50 of FIG. ordered. In other words, the control device 20 comprises a set of means configured as software (specific computer program product) and / or hardware (FPGA, PLD, ASIC, etc.) to implement the various steps. a control method 50 described hereinafter. FIG. 2 represents the main steps of a phase acquisition method 50 of initial acquisition of the Sun by the satellite 10, said satellite 10 having, immediately after separation with the launcher, a non-zero kinetic moment whose axis DEI can be estimated by means of suitable conventional sensors (gyroscope, stellar sensor, etc.). As illustrated in FIG. 2, the main steps of a control method 50 according to the invention are: - 51 actuator control of the satellite 10 so as to place said satellite 10 in an intermediate orientation, relative to at the kinetic moment, corresponding to an orientation of said satellite 10 in which the Y axis of rotation of the solar generators 12 is substantially orthogonal to the axis DH of the kinetic moment, - 52 control of the drive mechanism of the solar generators 12, prior to the less partially deployed, so as to orient said solar generators to the Sun. By "at least partially deployed solar generator" is meant that at least a portion of the photosensitive surface of said solar generator is both rotatable about the Y axis and available to generate electrical power. For example, if the photosensitive surface of a solar generator is distributed over several panels initially folded over each other, said solar generator can optionally be gradually deployed, by successively unfolding the different panels. After the satellite 10 has been placed in the intermediate orientation and after the solar generators 12, at least partially deployed, have been oriented towards the Sun, the control method 50 comprises steps of: - 53 command of the satellite actuators So as to reduce the kinetic momentum of said satellite 10, actuator control of the satellite 10 so as to place said satellite 10 in an acquisition orientation, with respect to the Sun, corresponding to an orientation of said satellite 10 in which the Y axis is substantially orthogonal to the direction Ds of the Sun relative to said satellite 10. Placement of the satellite in the intermediate orientation During step 51 of placing the satellite 10 in the intermediate orientation, it is not necessarily necessary to modify the kinetic moment of the satellite 10 in inertial reference, but to modify the orientation of said satellite 10 with respect to the axis DH of said moment kinét This makes the Y axis substantially orthogonal to the DH axis. It should be noted here that "substantially orthogonal" means that the setpoint angle 13 between the compared elements (the Y axis and the DH axis in the case above) is such that 1sin RI> 0.9. It should however be noted that the case where the setpoint angle p is such that 1sin RI = 1 (setpoint angle to obtain elements strictly orthogonal to errors) corresponds to a preferred embodiment. In the following description, in the following description, the term "substantially parallel" is understood to mean that the setpoint angle β between the compared elements is such that 'cos RI> 0.9, the case where the setpoint angle p is such that cos RI = 1 (setpoint angle to obtain elements strictly parallel to the errors) corresponding then to a preferred mode of implementation. It should be noted that the initial kinetic momentum (including the DH axis) is in principle inertial, that is to say invariant in the absence of external couples on the satellite 10. Thus, once the satellite 10 placed in the intermediate orientation, it is in principle not necessary to update said intermediate orientation. In the case where the initial kinetic moment would not be inertial, it is then advantageous, if the axis DH varies significantly, to update the intermediate orientation of the satellite 10 in order to follow the variations of the axis DH and maintain the Y axis substantially orthogonal to the axis DH of the kinetic moment.
    [0009] For example, if the Y axis corresponds to the minimum axis of inertia of the satellite 10, then the placement of the satellite 10 in the intermediate orientation is to align the maximum axis of inertia of said satellite 10 with the axis DH the kinetic moment. This corresponds, for the skilled person, to a forced passage in "flat spin" by damping the nutation of the satellite 10 relative to the axis DH of the kinetic moment. If necessary, the solar generators 12 may be at least partially deployed during or prior to the placement of the satellite 10 in the intermediate orientation, to ensure that the Y axis is the minimum axis of inertia of said satellite 10. From more generally, the Y axis is not necessarily the minimum axis of inertia of the satellite 10. For example, if the Y axis is the maximum axis of inertia, then the placement of the satellite 10 in the intermediate orientation consists in aligning the minimum axis of inertia of said satellite 10 with the axis DH of the kinetic moment, by an adapted control of the nutation of said satellite 10 with respect to the axis DH of said satellite 10.
    [0010] Such a change in the orientation of the satellite 10 with respect to the axis DH can be performed independently of the kinetic momentum, by any type of adapted actuator (chemical or electrical), including by inertial actuators whose capacity is lower than said cinematic moment.
    [0011] In a preferred embodiment, the placement of the satellite 10 in the intermediate orientation is effected by means of inertial actuators of said satellite 10, such as gyroscopic actuators and / or reaction wheels of said satellite 10. The control of said Inertial actuators, for controlling the nutation of the satellite 10 (alignment of the axis DH with the maximum axis of inertia or with the minimum axis of inertia of said satellite 10), may implement methods known to man of the art, for example the method described in US6382565. Orientation of Solar Generators When the satellite 10 is placed in the intermediate orientation, the solar generators 12, previously at least partially deployed, are oriented towards the Sun. Here, the term "oriented towards the Sun" means that said solar generators 12 are placed in respective orientations making it possible to optimize the sunlight of the photosensitive surfaces of said solar generators 12 over the duration of a complete rotation of the satellite 10 on itself . For example, the at least partially deployed solar generators 12 are placed in respective orientations making it possible to ensure that the average insolation of the photosensitive surfaces of said solar generators 12 over the duration of a complete rotation of the satellite on itself is greater than a predefined threshold value. As indicated above, in the intermediate orientation, the Y axis of rotation of the solar generators 12 is substantially orthogonal to the axis DH of the kinetic moment. Such orientation of the satellite 10 is particularly advantageous in that, although the attitude of the satellite 10 is not stabilized along three axes, it is nevertheless still possible to find respective orientations of the solar generators 12 to ensure autonomy electric satellite 10, as described below. The respective orientations of the solar generators 12 are, for example, determined as a function of the angle between the direction Ds of the Sun and the axis DH of the kinetic moment, called the "misalignment angle θ". The orientations of the solar generators 12 are for example kept fixed as long as the axis DH of the kinetic moment does not vary, which makes it possible to have a simple control of the solar generators 12 and to limit the use of the drive mechanisms as long as the kinetic moment of the satellite has not been reduced. Nothing, however, excludes, according to other examples, to vary the respective orientations of the solar generators 12 over time, even if the DH axis does not vary, to take account of the rotation of the satellite 10, for example to maximize the instantaneous sunshine of the photosensitive surface of at least one solar generator 12. In the remainder of the description, one places oneself in a nonlimiting manner in the case where the orientation of the solar generators 12 is kept fixed as long as the DH axis of kinetic moment does not vary.
    [0012] The term "pointing direction" of the photosensitive surface of a solar generator 12, a vector normal to said photosensitive surface, oriented side of said solar generator 12 which should be the Sun to generate electrical power. In a particular mode of implementation, the solar generators are maintained, during the kinetic moment reduction, so that the pointing direction of the photosensitive surface of each solar generator 12 is: - substantially orthogonal to the DH axis the kinetic momentum (in other words, said photosensitive surface is substantially parallel to the axis DH) when 'cos 01 <Vs1, Vs1 being a predefined threshold value, preferably between 0.3 and 0.6, - substantially parallel to the DH axis (in other words, said photosensitive surface is substantially orthogonal to the DH axis) when 'cos 01> Vs1, said photosensitive surface being oriented towards the Sun. FIG. 3 represents the average insolation obtained, as a function of the misalignment angle θ, over the duration of a complete rotation of the satellite 10 on itself, considering the threshold value Vs1 equal to approximately 0.34 and considering the Y axis and the DH axis strictly orthogonal.
    [0013] With such a control of the solar generators 12 when the satellite 10 is in the intermediate orientation, the average irradiation of the photosensitive surfaces of the solar generators 12 depends only on the misalignment angle θ.
    [0014] In particular, when cos 01> Vs1, the average irradiation of said photosensitive surfaces varies in cos 01 (each photosensitive surface being substantially orthogonal to the axis DH, the misalignment angle θ then corresponds to the distance from the angle of incidence of the Sun's rays on the photosensitive surfaces compared to the normal incidence). In addition, it can be seen that the average insolation of the photosensitive surfaces of the solar generators 12 is always greater than 30%. The average insolation of the photosensitive surfaces of the solar generators 12 will most often be lower than the optimum sunshine that can be obtained when the satellite 10 is in the acquisition orientation. However, such average sunshine is sufficient to ensure the electric autonomy of the satellite 10 in the sense that it is always possible to recharge the batteries of said satellite 10 by disabling if necessary the equipment that consumes the most electrical power.
    [0015] In a preferred embodiment, said solar generators 12 are maintained, during the reduction of the kinetic momentum, so that the respective photosensitive surfaces of said solar generators 12 are substantially parallel to the axis DH of the moment kinetic, and with respective pointing directions opposite when cos 01 <Vs2, Vs2 being a predefined threshold value. The threshold value Vs2 is preferably between 0.1 and 0.6, or even between 0.3 and 0.6. In particular modes of implementation, the threshold values Vs1 and Vs2 are equal. Considering for example that the axis DH of the kinetic moment is orthogonal to the direction Ds of the Sun (lcos 01 = 0), because the respective pointing directions of the photosensitive surfaces are opposite, there is always a photosensitive surface oriented on the side. of the Sun and a photosensitive surface oriented away from the Sun. If, on the contrary, the photosensitive surfaces of the solar generators 12 were oriented on the same side of the satellite 10, the two photosensitive surfaces would be simultaneously: - both oriented towards the Sun, during half the duration of the complete rotation of the satellite 10 on it themselves, both facing away from the sun for half the duration of said complete rotation.
    [0016] Such provisions therefore make it possible to limit the fluctuations of the instantaneous sunshine of the solar generators 12 around the average sunshine over the duration of a complete rotation of the satellite 10 on itself.
    [0017] It should be noted that the deployment of the solar generators 12 can be performed before, during or after the placement of the satellite 10 in the intermediate orientation. Similarly, the orientation of the solar generators 12 can be performed, if it depends only on the misalignment angle θ, after the DH axis has been estimated, before, during or after the placement of the satellite 10 in the intermediate orientation. As a result, the solar generators 12 can begin to supply electrical power before having begun to reduce the kinetic momentum of the satellite 10. Due to the intermediate orientation of the satellite 10 (in which the Y axis is substantially orthogonal to the DH axis), it is possible to ensure average sunning of the sensitive surfaces of the solar generators 12 by at least 30% over the duration of a complete rotation of the satellite 10 on itself. Such average sunshine is sufficient to ensure the electrical autonomy of the satellite 10 over time, and throughout the duration of the initial acquisition phase of the Sun. Therefore, it is possible to implement electric actuators, including electric thrusters, to then reduce the kinetic moment of the satellite 10. Where appropriate, the electric actuators can be activated discontinuously, in order to be able to recharge the batteries of the satellite 10 between two successive activations of said electric actuators, when said electric actuators are deactivated. This is made possible by a deployment of the solar generators 12 before reducing the kinetic moment of the satellite 10. The solar generators 12 can be fully deployed, or only partially deployed, especially if the initial kinetic moment is very important.
    [0018] In the case where the solar generators 12 are only partially deployed, they can for example be deployed progressively, as the kinetic moment of the satellite 10 decreases, or fully deployed after the kinetic moment of the satellite 10 has been reduced. If the initial kinetic moment is too great, it may be envisaged to perform a partial reduction of the kinetic momentum before deploying the solar generators 12 (where appropriate within the capacity of the batteries of the satellite 10). Most of the reduction in kinetic momentum, however, continues after at least partial deployment of said solar generators 12, the electrical autonomy of said satellite 10 then being ensured. Reduction of the Kinetic Moment of the Satellite Next, actuators of the satellite 10 are used to reduce the kinetic momentum of the satellite 10. By "reducing the kinetic momentum" of the satellite 10, it is intended to reduce the modulus of said kinetic momentum, preferably up to to reach a predefined threshold value. The reduction of the kinetic moment of the satellite 10 is preferably carried out by maintaining said satellite 10 in the intermediate orientation, that is to say by keeping the Y axis substantially orthogonal to the axis of the kinetic moment. The control of the satellite actuators 10, to reduce the kinetic moment of the satellite 10, can be performed in a conventional manner. As indicated above, the electrical autonomy of the satellite 10 is ensured by the placement of said satellite 10 in the intermediate orientation and by the at least partial deployment of the solar generators 12.
    [0019] In a preferred mode of implementation, the reduction of the kinetic momentum of the satellite 10, placed in the intermediate orientation, implements electrical thrusters (plasma) of said satellite 10. Indeed, the electric autonomy of the satellite 10 being ensured , the electric thrusters can be implemented despite the aforementioned drawbacks (unavailability the first days, low torque capacity, high power consumption). Nothing, however, excludes the use of other actuators of the satellite 10 to reduce the kinetic moment of the satellite 10, in addition or alternatively to electric thrusters. For example, in the case of a satellite 10 in LEO orbit, it is possible to implement magneto-couplers. According to another example, it is possible to implement chemical thrusters, which may have a lower torque capacity than those used in the prior art insofar as, the electrical autonomy of the satellite 10 being ensured, it it is no longer necessary to quickly place the satellite 10 in the acquisition direction. The kinetic momentum reduction is preferably performed without modifying the DH axis of the kinetic moment. Nothing, however, excludes also to change the DH axis, for example to start placing the satellite 10 in the acquisition direction. In such a case, it is then advantageous, if the DH axis varies significantly, to update the intermediate orientation of the satellite 10 in order to follow the variations of the DH axis and, if necessary, the orientation of the generators. solar 12.
    [0020] Placement of the satellite in the acquisition orientation Simultaneously with and / or after the reduction of the kinetic momentum, the satellite 10 is placed in the acquisition orientation, with respect to the Sun, in which the Y axis of the solar generators 12 is substantially orthogonal to the direction Ds of the Sun with respect to the satellite 10.
    [0021] In the remainder of the description, one places oneself in a nonlimiting manner in the case where the placement of the satellite 10 in the acquisition direction corresponds to a stabilization of the attitude of the satellite 10 along three axes. However, according to other examples, nothing prevents the acquisition of a non-zero rotation speed of the satellite 10 in the acquisition orientation. If necessary, the axis DH of the kinetic moment is modified so as to be made substantially parallel to the direction Ds of the Sun relative to the satellite 10 while maintaining the Y axis substantially orthogonal to the axis DH, so that the Y axis is substantially orthogonal to the direction Ds of the Sun throughout the duration of the rotation of the satellite 10 on itself.
    [0022] The placement of the satellite 10 in the acquisition orientation can implement any type of suitable actuator. The electrical autonomy of the satellite 10 over time being ensured, the placement of the satellite 10 in the acquisition orientation preferably implements electric actuators. The actuators used may be the same as those used to reduce the kinetic moment of the satellite 10, or other actuators. In a preferred mode of implementation, the placement of the satellite 10 in the acquisition orientation uses inertial actuators of said satellite 10, preferably gyroscopic actuators and / or reaction wheels of the satellite 10. The control of Inertial actuators of the satellite 10, for placing said satellite 10 in the acquisition orientation, can be carried out conventionally. For example, the inertial actuators can be implemented when the kinetic moment modulus of said satellite 10 becomes less than a predefined threshold value equal to or less than the capacity of said inertial actuators. When the satellite 10 is in the acquisition orientation, the solar generators 12 are preferably oriented so that the sun's rays have a substantially normal incidence on the photosensitive surfaces of the solar generators, in order to maximize the electrical power generated. FIG. 4 schematically represents a nonlimiting example of implementation of an initial acquisition phase of the sun 60 according to the invention: at the instant T1: the satellite 10 has just been separated from the launcher with a kinetic moment inertial equal to Ho.vo, where vo is a unit vector of the DH axis at time T1 and Ho is the kinetic moment modulus at time T1, at time T2: the kinetic moment Ho .vo was estimated by means of sensors of the satellite 10, - at time T3: the satellite 10 was placed in the intermediate orientation, with a kinetic moment always equal to Ho.vo, - at time T4: the solar generators 12 of the satellite 10 have been deployed (fully deployed in the nonlimiting example illustrated in FIG. 4); at the moment T5: the solar generators 12 have been oriented towards the sun 60, the photosensitive surfaces being, in the example shown, substantially parallel to the DH axis, at time T6: the kinetic momentum has been reduced to constant DH axis, so that the kinetic momentum is equal to Hi.vo, expression in which H1 <Ho is the kinetic momentum modulus at instant T6, at time T7: the satellite 10 has been placed in the acquisition orientation while continuing to reduce the kinetic momentum, the axis DH now substantially passes through the sun 60 and the kinetic momentum is equal to H2.v1, where H2 <H1 is the kinetic moment modulus at time T7 and v1 is a unit vector of axis DEI at time T7. More generally, it should be noted that the modes of implementation and realization considered above have been described by way of non-limiting examples, and that other variants are therefore possible. In particular, the invention has been described by considering an initial acquisition phase of the Sun. As indicated above, the invention is also applicable to other phases of acquisition of the Sun. In particular, the invention is applicable to the acquisition phase of the Sun in survival mode, for which the above observations remain valid, except that the solar generators 12 of the satellite 10 are then already deployed before the start of the phase acquisition of the Sun in survival mode. In addition, an acquisition phase control method 50 according to the invention can be combined with other control methods. In particular, a control method 50 according to the invention can be implemented only if the initial kinetic moment is very important. Thus, considering that only electrical actuators are implemented, the control of the acquisition phase can be adapted as a function of the initial kinetic momentum: if the initial kinetic moment is in the capacity of the inertial actuators of the satellite 10: Inertial actuators are implemented to reduce the kinetic moment of said satellite 10, - if the initial kinetic moment is not in the capacity of the inertial actuators, but is in the joint capacity of the electric thrusters and the batteries of the satellite 10: the thrusters electrical devices are implemented to reduce the kinetic moment of said satellite 10, - if the initial kinetic moment is not in the capacity of the inertial actuators, and is not in the joint capacity of the electric thrusters and the batteries of the satellite 10: a control method 50 according to the invention is then executed. The above description clearly illustrates that by its different characteristics and advantages, the present invention achieves the objectives it has set for itself. In particular, because the electrical autonomy of the satellite 10 is ensured by placing said satellite 10 in the intermediate orientation and by suitably orienting the solar generators 12, the acquisition phase of the Sun can only implement electric actuators. Therefore, the proposed solution is applicable in particular in the case of a satellite devoid of chemical thruster.
    权利要求:
    Claims (16)
    [0001]
    CLAIMS1 - Method (50) for controlling a phase of acquisition of the Sun by a spacecraft (10) with a non-zero kinetic momentum of axis DH, said spacecraft comprising a body, a solar generator (12) and a mechanism for driving said solar generator in rotation around a Y axis, characterized in that it comprises steps of: - (51) actuating actuators of the spacecraft (10) so as to place said spacecraft in an intermediate orientation, with respect to the kinetic moment, in which the Y axis is substantially orthogonal to the axis DH of the kinetic momentum, - (52) controls the driving mechanism of the solar generator (12), before at least partially deployed so as to orient said solar generator towards the Sun, - (53) actuating actuators of the spacecraft (10) so as to reduce the kinetic moment of said spacecraft, - (54) controlling actuators of the spacecraft. spacecraft (10) so as to place said craft in a direction of acquisition, with respect to the Sun, in which the Y axis is substantially orthogonal to the direction Ds of the Sun relative to the spacecraft.
    [0002]
    2 - Method (50) according to claim 1, characterized in that the placement of the spacecraft (10) in the intermediate orientation implements inertial actuators of said spacecraft, preferably gyroscopic actuators and / or wheels of reaction.
    [0003]
    3 - Method (50) according to one of claims 1 to 2, characterized in that the placement of the spacecraft (10) in the acquisition orientation uses inertial actuators of said spacecraft, preferably gyroscopic actuators and / or reaction wheels.
    [0004]
    4 - Process (50) according to claim 3, characterized in that the inertial actuators are implemented to place the spacecraft (10) in the acquisition orientation, when the kinetic moment of said spacecraft is less than a predefined threshold value.
    [0005]
    5 - Process (50) according to one of the preceding claims, characterized in that the reduction of the kinetic moment of the spacecraft (10), placed in the intermediate orientation, uses electric thrusters of said spacecraft.
    [0006]
    6 - Process (50) according to one of the preceding claims, characterized in that only electrical actuators are implemented during the acquisition phase of the Sun.
    [0007]
    7 - Method (50) according to one of the preceding claims, characterized in that, the maximum axis of inertia of the spacecraft (10), when the solar generator of said spacecraft is deployed, being orthogonal to the Y-axis, the solar generator (12) is at least partially deployed prior to the placement of the spacecraft (10) in the intermediate orientation and, during the placement of said spacecraft in said intermediate orientation, performs a nutation damping .
    [0008]
    8 - Method (50) according to one of claims 1 to 6, characterized in that, the minimum axis of inertia of the spacecraft (10) being orthogonal to the axis Y, is carried out, during the placing said spacecraft in said intermediate orientation, a nutation control for aligning the minimum axis of inertia of said satellite with the axis DH of the kinetic moment.
    [0009]
    9 - Method (50) according to one of the preceding claims, characterized in that the orientation of the solar generator (12) during the reduction of the kinetic moment is determined as a function of the angle between the direction Ds of the Sun and the axis DH of the kinetic moment, called "angle of offset e".
    [0010]
    10 - Process (50) according to claim 9, characterized in that the orientation of the solar generator (12) is maintained fixed relative to the body of the spacecraft (10) during the reduction of the kinetic moment.
    [0011]
    11 - Method (50) according to claim 10, characterized in that the solar generator (12) is maintained, during the kinetic moment reduction, so that a photosensitive surface of said solar generator is: - substantially parallel to the the axis DH of the angular momentum when cos 01 <Vs1, Vs1 being a predefined threshold value, - substantially orthogonal to the axis DH when 'cos 01> Vs1, said photosensitive surface being oriented towards the Sun.
    [0012]
    12 - Method (50) according to one of claims 10 to 11, characterized in that, the spacecraft (10) comprising two solar generators (12) mounted on respective drive mechanisms adapted to rotate said solar generators around respective parallel Y axes, said solar generators are maintained, during the kinetic momentum reduction, so that respective photosensitive surfaces of said solar generators are substantially parallel to the kinetic moment DH axis and with pointing directions respectively opposite when cos 6i <Vs2, Vs2 being a predefined threshold value.
    [0013]
    13 - Method (50) according to one of the preceding claims, characterized in that the acquisition phase of the Sun by the spacecraft (10) is controlled remotely by a device (20) for controlling, control signals being successively determined and sent to the spacecraft by said control device.
    [0014]
    14 - Computer program product characterized in that it comprises a set of program code instructions which, when executed by a processor, implement a method (50) for controlling a phase of a program. Sun acquisition by a spacecraft (10) according to one of the preceding claims.
    [0015]
    15 - Device (20) for controlling a phase of acquisition of the Sun by a spacecraft (10) with a non-zero kinetic moment of axis DH, said spacecraft comprising a body, actuators, a solar generator (12) ) and a drive mechanism of said solar generator in rotation about a Y axis, characterized in that it comprises means configured to control said spacecraft (10) according to one of claims 1 to 13.
    [0016]
    16 - Space system comprising a spacecraft (10) comprising a body, actuators, a solar generator (12) and a drive mechanism of said solar generator rotating about a Y axis, characterized in that it comprises a control device for a phase of acquisition of the Sun by said spacecraft (10) according to claim 15.
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    同族专利:
    公开号 | 公开日
    CN105899430B|2018-03-23|
    WO2015075237A1|2015-05-28|
    US20170183108A1|2017-06-29|
    CN105899430A|2016-08-24|
    EP3074309A1|2016-10-05|
    EP3074309B1|2017-07-05|
    US9988162B2|2018-06-05|
    FR3013685B1|2017-05-19|
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    US5535965A|1991-09-06|1996-07-16|Deutsche Aerospace Ag|Three-axis stabilized, earth-oriented satellite and a corresponding sun and earth acquisition process|
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    EP0788045A1|1996-02-05|1997-08-06|HE HOLDINGS, INC. dba HUGHES ELECTRONICS|Sun/earth acquisition without thrusters|
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    FR2980176A1|2011-09-19|2013-03-22|Astrium Sas|SATELLITE ATTITUDE CONTROL METHOD AND ATTITUDE CONTROL SATELLITE|
    FR2993863B1|2012-07-27|2014-08-29|Thales Sa|METHOD FOR REDUCING THE KINETIC MOMENT AND ATTITUDE CONTROL OF A SPATIAL MACHINE|FR3034535B1|2015-03-31|2018-08-17|Airbus Defence And Space Sas|METHOD AND DEVICE FOR CONTROLLING THE ATTITUDE OF A SPACE DEVICE|
    CN110712769B|2019-09-23|2021-02-05|北京控制工程研究所|Gyro-free sun orientation control method based on CMG|
    CN110641741B|2019-10-23|2020-12-18|北京国电高科科技有限公司|Double-freedom-degree solar panel control method and control system thereof|
    CN110963087B|2019-11-11|2021-10-29|上海航天控制技术研究所|Multi-mode complex-process control method for space station solar wing|
    CN111891396B|2020-08-12|2021-12-24|中国科学院微小卫星创新研究院|Small geostationary orbit satellite orbit transfer method and system|
    CN112777003B|2020-12-31|2021-09-21|中国人民解放军63923部队|Method and device for preventing damage of SADA arc discharge to satellite|
    法律状态:
    2015-11-23| PLFP| Fee payment|Year of fee payment: 3 |
    2016-11-29| PLFP| Fee payment|Year of fee payment: 4 |
    2017-04-07| CA| Change of address|Effective date: 20170301 |
    2017-04-07| CD| Change of name or company name|Owner name: AIRBUS DEFENCE AND SPACE SAS, FR Effective date: 20170301 |
    2017-11-30| PLFP| Fee payment|Year of fee payment: 5 |
    2019-11-29| PLFP| Fee payment|Year of fee payment: 7 |
    2020-11-27| PLFP| Fee payment|Year of fee payment: 8 |
    2021-11-30| PLFP| Fee payment|Year of fee payment: 9 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1361581A|FR3013685B1|2013-11-25|2013-11-25|METHOD AND DEVICE FOR CONTROLLING A SUN ACQUISITION PHASE BY A SPATIAL DEVICE|FR1361581A| FR3013685B1|2013-11-25|2013-11-25|METHOD AND DEVICE FOR CONTROLLING A SUN ACQUISITION PHASE BY A SPATIAL DEVICE|
    US15/039,018| US9988162B2|2013-11-25|2014-11-24|Method and device for control of a sunlight acquisition phase of a spacecraft|
    EP14802076.1A| EP3074309B1|2013-11-25|2014-11-24|Method and device for control of a sunlight acquisition phase of a spacecraft|
    PCT/EP2014/075419| WO2015075237A1|2013-11-25|2014-11-24|Method and device for control of a sunlight acquisition phase of a spacecraft|
    CN201480072346.0A| CN105899430B|2013-11-25|2014-11-24|For the method and apparatus for the daylight acquisition phase for controlling spacecraft|
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