• 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 present invention relates to a method (50) of attitude control of a spacecraft (10) rotating on itself with a total kinetic moment H TOT nonzero, said spacecraft (10) comprising a set of flywheels inertia (20) adapted to form an internal kinetic moment HACT. The attitude control method (50) comprises a step (52) of aligning the axis of said total kinetic moment HTOT with a main axis of inertia of the spacecraft (10), during which the flywheels of inertia (20) are controlled so as to form an internal kinetic moment HACT such that the following expression, in which J is the inertia matrix of the spacecraft (10): HACT X J-1 (HTOT ( X) J-1HTOT) - is negative if said main axis of inertia is the maximum axis of inertia of the spacecraft (10), - is positive if said main axis of inertia is the axis of minimal inertia of the spacecraft (10).
    公开号:FR3034535A1
    申请号:FR1552736
    申请日:2015-03-31
    公开日:2016-10-07
    发明作者:Nicolas Cuilleron；Philippe Laurens；Valerio Moro
    申请人:Airbus Defence and Space SAS；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD The present invention belongs to the field of attitude control of spacecraft, such as satellites, and more particularly relates to a method and a system for controlling the attitude of a spacecraft in rotation with itself with a total initial non-zero kinetic moment. By "attitude control" is meant here more particularly to modify the orientation of the spacecraft relative to the axis of said initial total kinetic moment, called "kinetic axis", that is to say to align, in a reference associated with the geometry of the spacecraft, said "gear mark", said kinetic axis with a predetermined axis in machine reference. STATE OF THE ART To modify the orientation of a spacecraft with respect to the kinetic axis, it is known to: - stop the rotation of the spacecraft in an inertial frame by absorbing said total kinetic moment by means of flywheels of inertia (reaction wheels, gyroscopic actuators), - of placing, by means of said flywheels, the spacecraft in the selected orientation with respect to the axis of said total kinetic moment, - of transferring kinetic moment total stored in flywheels to the spacecraft in the chosen orientation. Such an approach assumes, however, that the initial total kinetic momentum is in the absorption capacity of the flywheels, which can not always be guaranteed. In particular, the initial total kinetic moment, transferred to a satellite at the moment of separation with a launcher of said satellite, is generally too large to be absorbed by the flywheels of said satellite. For example, the initial total kinetic moment of a satellite after launch can be of the order of 500 Nms to 1000 Nms, whereas the capacity of the inertia flywheels embedded in a satellite is generally of the order of 50 Nms to 100 Nms In addition, even if the flywheels are dimensioned so as to have a capacity of 1000 Nms, particularly unfavorable conditions of separation could still transfer to the satellite an initial kinetic moment greater than 1000 3034535 2 Nms, outside the capacity absorption of flywheels. This is why, nowadays, the modification of the orientation of a satellite, with respect to the kinetic axis after launching, is generally preceded by a decrease in said total kinetic moment by means of 5 chemical thrusters, making it possible to restoring said total kinetic momentum in the absorption capacity of the flywheels. It is currently envisaged that future satellites will no longer be equipped with chemical thrusters, but only with electric (plasma) thrusters. However, electric thrusters can not be substituted for chemical thrusters, to reduce the total kinetic momentum of a satellite after separation, as long as the electric autonomy of said satellite is not ensured. However, the electric autonomy of the satellite can be ensured only by placing said satellite in a suitable orientation. DISCLOSURE OF THE INVENTION The present invention aims to remedy all or part of the limitations of the solutions of the prior art, in particular those set out above, by proposing a solution that makes it possible to modify the orientation of a machine the kinetic axis only by means of flywheels, even when said initial total kinetic moment is greater than the absorption capacity of the flywheels. For this purpose, and according to a first aspect, the invention relates to a method of attitude control of a spacecraft in rotation on itself with a total kinetic momentum 11ToT nonzero, said spacecraft comprising a set of flywheels. inertia adapted to form an internal kinetic moment 25 HACT of any axis in a reference machine. The attitude control method comprises a step of aligning the axis of said total kinetic moment HTOT with a main axis of inertia of the spacecraft, during which the flywheels are controlled so as to form an internal kinetic moment HAcT such that the following expression: HACT X J-1 (HTOT 0 r1HTOT) is negative for the duration of the alignment step with the main axis of inertia if said main axis of target inertia is the maximum axis of inertia of the spacecraft, 3034535 3 - is positive throughout the duration of the alignment step with the main axis of inertia if the said principal axis of inertia is the minimum axis of inertia of the spacecraft, expression in which J is the inertial matrix of the spacecraft in machine reference 5, the operator x is the scalar product between two vectors and the operator 0 is the vector product between two vectors. Indeed, the inventors have found that such provisions always make it possible to align the axis of the total kinetic moment with the main axis of inertia aimed at (ie the maximum axis of inertia or the minimum axis of inertia). ). In the case where the principal axis of inertia is the maximum axis of inertia, such provisions also make it possible to significantly accelerate the convergence compared to the case where no control is performed (a spacecraft subject to a total non-zero kinetic moment having, because of internal energy dissipations and in the absence of control, a tendency to converge slowly towards an orientation in which the axis of said total kinetic moment is aligned with the axis of inertia maximum). In addition, the sign of the expression HAcr X J-1 (1TOT ° J-111TOT) is entirely determined by the direction and direction of the internal kinetic moment HAcT, and is independent of the modulus of said internal kinetic moment HAcT. Consequently, the attitude control method which is the subject of the invention can be implemented even when the initial total kinetic moment Hun 'is greater than the capacity of absorption of the flywheels of the spacecraft. In other words, it can always be ensured that the internal kinetic moment HAcT respects the maximum kinetic momentum storage capacity of the flywheels and the maximum capacity of torque formation of the flywheels. In particular embodiments, the attitude control method may further comprise one or more of the following characteristics, taken in isolation or in any technically possible combination.
    [0002] In particular embodiments, the flywheels are controlled so as to form an internal kinetic moment HAcT for which an angle θ between said internal kinetic moment HAcT and the vector 3034535 4 1-1 (1-1Tcyr 0 1-11froT) checks, during the entire alignment step with the main axis of inertia, the following expression: Icosel> 0.9 Such provisions make it possible to significantly accelerate convergence towards the main axis of inertia referred to, it being understood that the direction of the vector 1-1 (FIT0T 0 1-111T0T) is the optimal direction to have the fastest convergence to the main axis of inertia referred. In particular embodiments, the flywheels are controlled during the entire alignment step with the main axis of inertia, so as to form an internal kinetic moment HAcT: HACT = Kv - U expression in which Kv is a negative scalar parameter if the principal axis of inertia is the maximum axis of inertia of the spacecraft or positive if the principal axis of inertia is the axis of minimal inertia of said spacecraft, and U corresponds to the unit vector: III 1 (Hurr 01 1HT0T) 11 In particular embodiments, the attitude control method comprises, after the step of alignment with the main axis of inertia, an alignment step with a predetermined axis X in machine mark, during which the flywheels are controlled so as to slave the components of the internal kinetic moment HAcT along Y, Z, 20 axes transverse to the X axis, on instructions respectively hy and hz d determined according to the components of the speed of rotation of the spacecraft along said axes Y, Z: the setpoint hy of the internal kinetic moment HAcT along the Y axis is determined according to a proportional-type control law integral to from the component r of the rotational speed along the Z axis, the setpoint hz of the internal kinetic moment HAcT along the Z axis is determined according to a proportional-integral control law from the component q of the speed of rotation = I 1 (11ToT 0 i 1Hr: 1a) U 3034535 5 along the Y axis. Such provisions make it possible to reduce the possible nutation of the spacecraft around the main axis of inertia targeted. Indeed, the control law of the flywheels during the alignment step with the main axis of inertia is non-linear makes it possible to align the kinetic axis with the maximum axis of inertia or with the minimum axis of inertia regardless of the kinetic moment and torque formation capacity of the spacecraft flywheels. However, the vector 1-1 (Flpyr 0 IlifroT) tends to zero at the approach of the principal axis of inertia aimed, so that any error on the knowledge of the inertia matrix J can result in residual nutation around said main axis of inertia. The linear proportional-integral control law makes it possible to ensure convergence towards an axis X, close to (within the limit of the capacity of flywheels) or coincides with the main axis of inertia aimed at, and damping the nutation to a zero value around said axis X. In particular embodiments, the hy and hz setpoints of the internal kinetic moment HAcT along the Y and Z axes are respectively connected to the components respectively r and q of the rotational speed of the spacecraft by the following transfer functions, expressed in the Laplace domain: {hy = Kz (1 + (j ')) r wS y hz = -Ky (1 + - s) q expressions in which: - s is the Laplace variable, - Ky and Kz are scalar parameters of the same sign constant over time, 25 - (1) Y and wz are constant positive scalar parameters over time . In particular embodiments, the step of alignment with the X axis is performed at constant total angular momentum 11ToT in an inertial frame.
    [0003] In particular embodiments, the alignment step 3034535 6 with the main axis of inertia is performed at total kinetic moment HTOT constant in inertial reference. In particular embodiments, the attitude control method comprises, after the step of alignment with the main axis of inertia, a step of modifying the total kinetic moment HT0T in inertial reference by means of thrusters of the spacecraft. In particular modes of implementation, the alignment step with the main axis of inertia is performed before any modification of the total kinetic moment Hun, inertial reference.
    [0004] 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 an attitude control method according to the present invention. any of the embodiments of the invention.
    [0005] According to a third aspect, the invention relates to a system for controlling the attitude of a spacecraft rotating on itself with a total kinetic moment Hun, not zero, said spacecraft comprising a set of adapted flywheels to form an internal kinetic moment HAcT of any axis in a machine coordinate system, said attitude control system comprising means configured to implement an attitude control method according to any one of the implementation modes of the invention. PRESENTATION OF THE FIGURES The invention will be better understood on reading the following description, given by way of non-limiting example, and with reference to the figures which represent: FIG. 1: a schematic representation of a satellite comprising a Attitude control system, - Figure 2: a diagram illustrating the main steps of a particular embodiment of an attitude control method, - Figure 3: a diagram illustrating the main steps of a preferred embodiment of an attitude control method, FIG. 4: curves illustrating the performance of an attitude control method according to the invention. In these figures, identical references from one figure to another designate identical or similar elements. For the sake of clarity of the figures, the elements shown are not necessarily to scale, unless otherwise indicated. DETAILED DESCRIPTION OF EMBODIMENTS The present invention relates to the attitude control of a spacecraft rotating on itself with a total kinetic moment Hun 'initial non-zero. The axis of the total kinetic moment HTOT is hereinafter referred to as "kinetic axis". By "attitude control" is meant here modifying at least the orientation of the spacecraft relative to the kinetic axis, that is to say align, in a reference associated with the geometry of the machine spatial, said "gear mark", said kinetic axis with a predetermined axis in machine reference.
    [0006] In the remainder of the description, the case in which the spacecraft is a satellite 10 and in which the initial total momentum HTOT corresponds to the momentum transferred to said satellite 10 during the separation with a launcher of said satellite is considered in a nonlimiting manner. . For example, the satellite 10 has been placed into the Geostationary Transfer Orbit (GTO) by the launcher and is destined to perform its mission in GEO ("Geostationary Orbit") orbit. However, there is nothing to exclude, according to other examples, other types of spacecraft (space shuttle, spacecraft, etc.). In addition, the invention is more generally applicable to any spacecraft whose kinetic axis orientation is desired to be modified in the vehicle mark and whatever the current or final orbit of said spacecraft (LEO or " Low Earth Orbit ", MEO or" Medium Earth Orbit ", GTO, GEO, etc.). FIG. 1 represents a particular embodiment of a satellite 10 comprising an attitude control system. In practice, the attitude control system may also include other elements than those shown in FIG. 1, which are outside the scope of the invention. In the nonlimiting example illustrated in FIG. 1, the attitude control system comprises a set of flywheels 20 adapted to form an internal kinetic moment HAcT of any axis in machine reference.
    [0007] As illustrated in FIG. 1, the attitude control system comprises, for example, at least three flywheels 20, such as reaction wheels and / or gyroscopic actuators, adapted to form an internal kinetic moment. HAcT of any axis in machine reference. For example, the attitude control system has three respective linearly independent unit vector axis reaction wheels. The satellite attitude control system 10 also has a control device 30. The control device 30 controls the attitude of the satellite 10 and controls for this purpose, in particular, the flywheels 20.
    [0008] The control device 30 comprises, for example, at least one processor and at least one electronic memory in which a computer program product is stored, in the form of a set of program code instructions to be executed to implement the various steps of a satellite attitude control method 50. In a variant, the control device 30 also comprises one or more programmable logic circuits of the FPGA, PLD, etc. type, and / or integrated circuits. ASICs adapted to implement all or part of said steps of the attitude control method 50 of the satellite 10. In other words, the control device 30 comprises a set of software configured means 20 (program product specific computer) and / or hardware (FPGA, PLD, ASIC, etc.) to implement the various steps of an attitude control method 50 described below. In the example illustrated in FIG. 1, the control device 30 is embedded in the satellite 10. More generally, the control device 30 can be embedded in the satellite 10 or in one or more remote equipment of said satellite 10 , in particular terrestrial equipment (the flywheels 20 being optionally remote controlled). Nothing further excludes, according to other examples, having a control device 30 distributed between the satellite 10 and one or more other equipment 30 distant from said satellite 10. In the case where the control device 30 is at least partially embedded in an equipment remote from the satellite 10, said remote equipment and the satellite 10 comprise respective conventional means of remote communication.
    [0009] In principle, a method 50 of attitude control according to the invention mainly comprises a step 52 of aligning the axis of said total kinetic moment 11ToT with a main axis of inertia of the satellite 10, which puts a law non-linear control. The main axis of inertia referred to is either the axis of maximum inertia ("flat spin") of the satellite 10, or the minimum axis of inertia of said satellite 10. During the alignment step 52 with the main axis of inertia, the control device 30 controls the flywheels 20 so as to form an internal kinetic moment HAcT such that the following expression: HACT X J-1 (u 0 J-1HTOT) 10 is of constant sign throughout the duration of said alignment step 52 with the main axis of inertia. In the previous expression, J is the inertial matrix of the satellite 10 in machine reference, the operator x is the dot product between two vectors and the operator CD is the vector product between two vectors. More particularly, the flywheels 20 are controlled so that the above expression: - is negative throughout the duration of the alignment step 52 with the main axis of inertia if said main axis of inertia target is the maximum axis of inertia of the satellite 10, is positive throughout the duration of the alignment step 52 with the main axis of inertia if said main axis of inertia is the axis of the minimum inertia of the satellite 10. The sign of the expression HAcr X J-1 (HTOT litroT) is entirely determined by the direction and direction of the internal kinetic moment HAcT, and is independent of the modulus of said internal kinetic moment HAcT. Therefore, the attitude control method 50 may be implemented even when the initial total kinetic moment Hun 'is greater than the absorption capacity of the spacecraft flywheels. In other words, it can always be ensured that the internal kinetic moment HAcT, to be formed during the alignment step 52 with the main axis of inertia, respects the maximum capacity of kinetic moment formation and The attitude control system comprises, for example, a measuring device 303 (not shown in the figures) adapted to measure the inertial rotation speed of the satellite 10, from which the control device 30 can conventionally determine the total kinetic moment Hun, initial of the satellite 10.
    [0010] Then, the control device 30 determines, for example, the components of the total kinetic moment Hun, in the apparatus reference at the instant considered, and deduces the conditions that the internal kinetic moment HAcT to be formed must verify to ensure that the sign of the expression HAcr X ri (11TOT 0 1-11froT) is negative if the principal axis of inertia is the maximum axis of inertia, positive if the principal axis of inertia is the axis of minimal inertia. The control device 30 then determines a suitable setpoint of internal kinetic moment HAcT, and a corresponding torque command to be formed by said flywheels 20 to obtain said setpoint of the internal kinetic moment HAcT. The actual value of the internal kinetic moment HAcT is for example determined by means of measurements of the respective rotational speeds of the flywheels 20, and the difference between the real value and the internal kinetic moment value set HAcT is for example used, conventionally, to update the torque commands to be formed by said flywheels 20. These various steps are iterated over time to account for the rotation of the satellite 10 (and variations in the components of the total kinetic moment Hun, in the apparatus reference) to ensure that the sign of the expression HAcT X J-1 (1TOT 0 riIOTOT) remains constant throughout the duration of the step 52 of aligning the kinetic axis with the main axis of inertia.
    [0011] As previously indicated, the direction of the 1-1 vector (1-1Tcyr 01-11froT) is the optimum direction for having the fastest convergence to the main axis of inertia. In preferred embodiments, the flywheels 20 are thus controlled so as to form an internal kinetic moment HAcT 30 whose direction, throughout the duration of the step 52 of alignment with the main axis of Inertia, is not too far from that of the vector 1-1 (1-1Tcyr 01-11froT) - More particularly, the flywheels 20 are 303 4 53 5 11 controlled so that the absolute value of the cosine of an angle θ between said internal kinetic moment HAcT and the vector 1-1 (-1Tcyr 1-11froT) is, throughout the duration of the alignment step 52 with the main axis of inertia, greater than 0.9 ( lcosel> 0.9), or even higher than 0.98 (lcosel> 0.98).
    [0012] Preferably, the flywheels 20 are controlled, during the entire duration of the alignment step 52 with the main axis of inertia, so as to form an internal kinetic moment HAcT substantially of the same direction as that of the vector J 10-1ToT 0 J-11-1T0T) HACT = KvU expression in which Kv is a scalar parameter and U is the direction of the vector J (H-1 .-- ToT 01-11-1T0T) = I 1 (11ToT 111ToT) U III 1- (HT0T 1HT0T) 11 The scalar parameter Kv is either negative if the principal axis of inertia is the maximum axis of inertia of the satellite 10, or positive if said main axis of inertia target is the minimum axis of inertia of said satellite 10. The scalar parameter Kv is constant or variable over time, and its value is determined so as to ensure that the internal kinetic moment HAcT, to be formed in the course of time. step 52 of alignment with the main axis of inertia, always respects the maximum capacity of kinetic moment formation and For example, the value of the scalar parameter Kv verifies the following expressions: Tmax 'Kyi II (v1 Hmax 20 expressions in which: Tmax corresponds to the maximum capacity of torque formation of the flywheels of inertia 20, - Û corresponds to the time derivative of the direction U of the vector J-10-froT J-1 IITOT) 25 - Hmax corresponds to the maximum capacity for kinetic moment formation of the flywheels 20. The value of the parameter of the Kv scalar parameter can also be 3034535 12 limited when the total kinetic moment direction Hun, is approaching the main axis of inertia. For example, when the angle cp between the direction of the total kinetic moment Hun, and the principal axis of inertia becomes less than 15 °, it is possible to apply a factor simp / sin (15 °) which decreases progressively to As the angle cp decreases. FIG. 2 represents a preferred embodiment of a satellite attitude control method 50. As illustrated by FIG. 2, the illustrated attitude control method 50 comprises, after step 52 alignment with the main axis of inertia 10 referred (maximum axis of inertia or minimum axis of inertia), a step 54 of alignment of the kinetic axis with a predetermined axis X in vehicle mark, during which uses a control law different from that used during the step 52 of alignment with the main axis of inertia referred to, in this case a proportional-integral linear control law.
    [0013] More particularly, during step 54 of alignment with the X axis, the flywheels 20 are controlled so as to slave the components of the internal kinetic moment HAcT along Y, Z axes transverse to the X axis and forming therewith a particular gear mark, on respectively hy and hz setpoints determined as a function of the components of the inertial rotation speed of the satellite 10 along said Y axes, Z: - the target hy of the kinetic moment internal HAcT along the Y axis is determined according to a proportional-integral type control law from the component r of the speed of rotation along the Z axis, the setpoint hz of the internal kinetic moment HAcT along the Z axis is determined according to a proportional-integral control law from the component q of the speed of rotation along the Y axis.
    [0014] As illustrated in FIG. 2, the target axis X is preferably coincident with the principal axis of inertia, in which case the proportional-integral control law serves to dampen the nutation to a minimum. zero value 3034535 13 about said main axis of inertia referred to. However, the proportional-integral control law also makes it possible to ensure convergence, from a kinetic axis substantially aligned with the principal axis of inertia aimed at, towards an axis X distinct from said main axis of inertia. and damping the nutation to a zero value about said X axis. Where appropriate, however, the X axis must be close enough to the main axis of inertia so that alignment of the kinetic axis with said X axis can be performed by means of the flywheels 20. The X axis can be considered close enough to the main axis of inertia when the inertia products Ixy and Ixz with the X axis check the 10 following expressions: PX & II HTOT II G IIHmaxil Ix expressions in which Ix corresponds to the inertia along the X axis. However, it is preferable to provide an additional margin to ensure that the alignment of the kinetic axis with said axis X can actually be carried out at yen flywheels 20: 11xy I - IIFIT0T II <111-Imax11 Ix P IlxzlIIHT0T11 IIHmax11 - <Ix P 15 expressions in which p corresponds to a predefined factor greater than one (p> 1), for example equal to two ( p = 2). In particular embodiments, the instructions hy and hz of the internal kinetic moment HAcT along the axes Y and Z, respectively, are connected to the components respectively r and q of the speed of rotation of the satellite 10 by the following transfer functions, expressed in the Laplace domain: ## EQU2 ## where: the Laplace variable, - Ky and Kz are dimensionless scalar parameters of the same sign constant over time, - wy and wz are positive scalar parameters, homogeneous at 5 pulsations (s-1), constant over time . The use of the Laplace domain in the field of servo systems is quite conventional and considered as known to those skilled in the art. Furthermore, the adjustment of the scalar parameters KY, Kz, wy and wz, given the above expressions of the hy and hz setpoints of the internal kinetic moment HAcT, is considered within the abilities of those skilled in the art. As indicated above, the attitude control method 50 makes it possible to align the kinetic axis with a main axis of inertia (maximum axis of inertia or minimum axis of inertia) irrespective of the formation capacity of the axis of inertia. kinetic moment and torque of the flywheels 20 of the satellite 10.
    [0015] Therefore, step 52 of alignment of the kinetic axis with the main axis of inertia target can be performed before any modification of the total kinetic moment HTOT in inertial reference. Likewise, the step 54 of aligning the kinetic axis with an axis X coinciding with or near the principal axis of inertia aimed at, can also be performed before any modification of the total kinetic moment HTOT in inertial reference. However, according to other examples, there is nothing to preclude modifying the total kinetic moment HTOT before and / or during aligning the kinetic axis with a predetermined axis in machine reference. FIG. 3 shows a preferred embodiment of a satellite attitude control method 10. As illustrated in FIG. 3, the attitude control method 50 comprises, after step 52 alignment with the main axis of inertia referred to, a step 56 of changing the total kinetic moment HTOT in inertial reference by means of thrusters (not shown in the figures) of the satellite 10, controlled by the device 30 control. The step 56 of modifying the total kinetic moment HTOT by means of propellants of the satellite 10 is conventional and is considered as known to those skilled in the art. For example, step 56 of altering the total kinetic moment HTOT may comprise the reduction of said total kinetic moment HTOT until it is in the capacity of absorption of flywheels 20. In the example 2, the attitude control method 50 also comprises the step 54 of alignment with an axis X, which implements a linear control law to dampen the nutation, which is for example also performed before step 56 of modification of the total kinetic moment Hun, inertial reference. In the example illustrated in Figure 3, the X axis is considered merged with the main axis of inertia referred. Preferably, the thrusters used during step 56 of modifying the total kinetic moment HTOT are electric thrusters.
    [0016] Advantageously, the main axis of inertia referred to (or, where appropriate, the target X axis) is an axis according to which the satellite 10 can ensure its electric autonomy by means of solar generators (not shown in the figures), so that it is possible to use electric thrusters to change the total kinetic moment Hun, inertial reference.
    [0017] In the remainder of the description, reference is made, by way of example, to the case where the solar generators of the satellite 10 are rotatable about an axis of rotation substantially orthogonal to the axis of rotation. maximum inertia of said satellite 10 (this will generally be the case if the solar generators are arranged on either side of a body of said satellite 10). The orientation of said solar generators around the axis of rotation is controlled by means of drive mechanisms. In this case, the main axis of inertia referred to during the alignment step 52 (and, if applicable, the X axis referred to during the alignment step 54) corresponds to the axis of maximum inertia of said satellite 10. At the end of this step, the kinetic axis is therefore substantially orthogonal to the axis of rotation of the solar generators. It can be shown that it is then always possible to find an orientation of the solar generators to ensure that the average insolation of photosensitive surfaces of said solar generators, over the duration of a complete rotation of the satellite 10 on itself, is greater than a predefined threshold value 30, for example greater than 30%. By placing the solar generators in such an orientation, it is then possible to ensure the electrical autonomy of the satellite 10, so that electric thrusters can be implemented to change the total kinetic moment Hun 'of the satellite 10.
    [0018] The foregoing description clearly illustrates that by its different features and advantages, the present invention achieves the objectives it has set for itself. In particular, the invention makes it possible to modify the orientation of a satellite 10 with respect to a kinetic axis by means of only 5 flywheels 20, even when the initial total angular momentum 11ToT is greater than the capacity of absorption of said flywheels 20, thanks to a nonlinear control law. FIG. 4 shows simulation results illustrating the performance of an attitude control method 50 according to the invention during step 52 of aligning the axis of the total kinetic moment HTOT with the main axis of inertia of a satellite 10. The inertia of the satellite 10 considered are the following: - Ix = 70000 kg.m2, - ly = 10000 kg.m2, 15 - lz = 50000 kg.m2. so that the main axis of inertia of said satellite 10 is substantially the axis X. In addition, the maximum capacity Hmax kinetic moment of the flywheels 20 of the satellite 10 considered is equal to 25 Nms, and the maximum capacity Tmax in pairs is equal to 0.2 Nm
    [0019] The part a) of FIG. 4 represents the rotation speeds, Qx, l / y, fz (expressed in rad / s) of the satellite 10. It can be seen that the initial total angular momentum 11ToT is close to the Y axis. In addition, the initial total kinetic moment 11ToT considered is 250 N-ms, so that it is much greater than the maximum capacity Hmax in kinetic moment of the flywheels 20 of the satellite 10.
    [0020] Part b) of FIG. 4 represents the setpoints hx, hy, hz of the internal kinetic moment HAcT during the alignment step 52. As illustrated by part a) of FIG. 4, the axis of the total kinetic momentum 11ToT converges well towards the X axis of the satellite 10, and without the internal kinetic momentum HACT exceeding the maximum formation capacity Hmax of The kinetic moment of the flywheels 20 of the satellite 10, as shown in part b) of FIG.
    权利要求:
    Claims (10)
    [0001]
    CLAIMS 1 - A method (50) for controlling the attitude of a spacecraft (10) rotating on itself with a total angular momentum 11ToT, said spacecraft (10) comprising a set of flywheels (20). ) adapted to form an internal kinetic moment HAcT of any axis in an apparatus reference, characterized in that it comprises a step (52) for aligning the axis of said total kinetic momentum 11ToT with a main axis of inertia of the spacecraft (10), in which the flywheels (20) are controlled to form an internal kinetic moment HAcT such as the following expression: HACT X J-1 (u 0 J-1uTOT) - is negative throughout the duration of the alignment step (52) with the main axis of inertia if the said principal axis of inertia is the maximum axis of inertia of the spacecraft (10), - is positive during the entire duration of the alignment step (52) with the main axis of inertia if the said principal axis of inertia is xe of minimum inertia of the spacecraft (10), expression in which J is the inertia matrix of the spacecraft (10), the operator x is the dot product between two vectors and the operator CD is the vector product between two vectors.
    [0002]
    2 - Method (50) according to claim 1, wherein the flywheels (20) are controlled so as to form an internal kinetic moment HAcT for which an angle 0 between said internal kinetic moment HAcT and the vector J 1 (1froT rittroT) checks, during the entire duration of the step (52) of alignment with the main axis of inertia, the following expression: Icosel> 0.9
    [0003]
    3 - Process (50) according to claim 2, wherein the flywheels (20) are controlled during the entire duration of the step (52) of alignment with the main axis of inertia, so as to to form an internal kinetic moment HACT HACT = Kv - U expression in which Kv is a negative scalar parameter if the principal axis of inertia of interest is the maximum axis of inertia of the spacecraft (10) or positive if said main axis of inertia referred to is the minimum axis of inertia of said spacecraft, and U corresponds to the unit vector: = 1-1 (HT0T 01-111r0T) U 111-1 (-froT 01 -111T0T) II
    [0004]
    4 - Method (50) according to one of the preceding claims, comprising, after the step (52) of alignment with the main axis of inertia, a step (54) of alignment with a predetermined axis X in machine reference, during which the flywheels (20) are controlled so as to slave the components of the internal kinetic moment HAcT along Y, Z axes, transverse to the X axis, respectively on hy and hz instructions 10 determined according to the components of the speed of rotation of the spacecraft (10) along said Y, Z axes: the setpoint hy of the internal kinetic moment HAcT along the Y axis is determined according to a proportional-type control law; integral from the component r of the speed of rotation 15 along the axis Z, the setpoint hz of the internal kinetic moment HAcT along the axis Z is determined according to a proportional-integral type of control law from the component q the next rotation speed the Y axis. 20
    [0005]
    5 - Process (50) according to claim 4, wherein the hy and hz instructions of the internal kinetic moment HAcT along the Y and Z axes respectively are connected to the components respectively r and q of the speed of rotation by the following transfer functions, expressed in the Laplace domain: {hy = Kz (1 + -wsz) r hz = -Ky (1 + (1`7) q 25 expressions in which: - s is the Laplace variable, - Ky and Kz are scalar parameters of the same sign constant over time, 3034535 19 - wy and wz are constant positive scalar parameters over time.
    [0006]
    6 - Process (50) according to one of claims 4 to 5, wherein the step (54) of alignment with the X axis is performed at total kinetic moment HTOT 5 constant in inertial reference.
    [0007]
    7 - Method (50) according to one of the preceding claims, wherein the step (52) of alignment with the main axis of inertia is performed at total kinetic momentum constant 11ToT in inertial reference.
    [0008]
    8 - Method (50) according to one of the preceding claims, comprising, 10 after the step (52) of alignment with the main axis of inertia, a step (56) for modifying the total angular momentum 11ToT benchmark inertial using thrusters of the spacecraft (10).
    [0009]
    9 - Computer program product characterized in that it comprises a set of program code instructions which, when executed by a processor, implement an attitude control method (50) according to the present invention. one of the preceding claims.
    [0010]
    10- attitude control system of a spacecraft (10) in rotation on itself with a total angular momentum 11ToT nonzero, said spacecraft comprising a set of flywheels (20) adapted to form a 20 internal kinetic moment HAcT of any axis in a machine mark, characterized in that it comprises means configured to control the flywheels (20) according to one of claims 1 to 8.
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    同族专利:
    公开号 | 公开日
    JP6433607B2|2018-12-05|
    US20180072435A1|2018-03-15|
    FR3034535B1|2018-08-17|
    US10407186B2|2019-09-10|
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    JP2018511895A|2018-04-26|
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    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    EP2690020A2|2012-07-27|2014-01-29|Thales|Method for reducing the angular momentum and controlling the attitude of a spacecraft|
    US7185855B2|2004-04-30|2007-03-06|Honeywell International, Inc.|Method and system for steering a momentum control system|
    US7900874B2|2008-01-22|2011-03-08|Harvey Emanuel Fiala|Device to move an object back and forth|
    US8066226B2|2008-01-22|2011-11-29|Fiala Harvey E|Inertial propulsion device to move an object up and down|
    JP5197498B2|2009-06-05|2013-05-15|三菱電機株式会社|Gimbal control device|
    JP5484262B2|2010-08-31|2014-05-07|三菱電機株式会社|Spacecraft attitude control device|
    CN102810936B|2012-08-01|2015-05-13|北京工业大学|Peripherally-actuated auxiliary moment generator for under-actuated system|
    US10005569B2|2013-07-02|2018-06-26|University Of Florida Research Foundation, Inc.|Triple flywheel assembly with attitude jitter minimization|
    FR3013685B1|2013-11-25|2017-05-19|Astrium Sas|METHOD AND DEVICE FOR CONTROLLING A SUN ACQUISITION PHASE BY A SPATIAL DEVICE|US10005568B2|2015-11-13|2018-06-26|The Boeing Company|Energy efficient satellite maneuvering|
    CN106882398B|2017-02-15|2019-08-30|上海航天控制技术研究所|A kind of control method of attitude control thruster|
    CN107472396B|2017-09-26|2021-04-27|北京航空航天大学|Quadruped robot capable of realizing air posture adjustment|
    CN110356589B|2019-06-04|2021-01-05|宁波天擎航天科技有限公司|Method and device for controlling water attack prevention of multiplexed side jet system and computer equipment|
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    2016-10-07| PLSC| Publication of the preliminary search report|Effective date: 20161007 |
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    2018-03-30| PLFP| Fee payment|Year of fee payment: 4 |
    2019-04-01| PLFP| Fee payment|Year of fee payment: 5 |
    2019-08-09| RN| Application for restoration|Effective date: 20190703 |
    2019-08-16| FC| Decision of inpi director general to approve request for restoration|Effective date: 20190709 |
    2020-03-30| PLFP| Fee payment|Year of fee payment: 6 |
    2021-12-10| ST| Notification of lapse|Effective date: 20211105 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1552736|2015-03-31|
    FR1552736A|FR3034535B1|2015-03-31|2015-03-31|METHOD AND DEVICE FOR CONTROLLING THE ATTITUDE OF A SPACE DEVICE|FR1552736A| FR3034535B1|2015-03-31|2015-03-31|METHOD AND DEVICE FOR CONTROLLING THE ATTITUDE OF A SPACE DEVICE|
    JP2017557331A| JP6433607B2|2015-03-31|2016-03-31|Method and device for controlling the attitude of a spacecraft|
    EP16712898.2A| EP3248079B1|2015-03-31|2016-03-31|Method and device for controlling attitude of a spacecraft|
    PCT/EP2016/057051| WO2016156487A1|2015-03-31|2016-03-31|Method and device for controlling attitude of a spacecraft|
    CN201680020554.5A| CN107438806B|2015-03-31|2016-03-31|Method and device for controlling the attitude of a spacecraft|
    US15/563,271| US10407186B2|2015-03-31|2016-03-31|Method and device for controlling attitude of a spacecraft|
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