• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to a method for guiding the setting of a satellite. It comprises the following steps carried out during a predefined current cycle: A) determining on the ground, a vector orientation law pushed, and a history of state variables and satellite state variables variables for the transfer of a starting orbit towards a target orbit using the theory of optimal control, B) from the orientation law, and the historical determination on the ground a law of evolution of the rotation of the satellite around the thrust vector, C) representing, according to a predetermined format, the evolution of the state and adjoint state variables to obtain first parameters, D) representing, according to a predetermined format, a law of evolution of the rotation to obtain second parameters, E) concatenate the first and second parameters to obtain a satellite guide plane, F) download on board the guide plane, G) iterate periodically according to a predefined period less than the duration of the cycle of guidance: - g1) reconstruct a guidance set on board the satellite, - g2) execute on board the satellite the setpoint by applying a closed control loop, H) measure on the ground the real orbital trajectory of the satellite, I) repeat the steps A) to H) with the trajectory measured at the end of the cycle as starting orbit of the next cycle, until reaching the target orbit.
    公开号:FR3030456A1
    申请号:FR1402879
    申请日:2014-12-17
    公开日:2016-06-24
    发明作者:Joel Amalric；Thierry Dargent；Bris Christophe Le
    申请人:Thales SA；
    IPC主号:
    专利说明:

    [0001] The present invention relates to the positioning or the transfer of orbit for satellites equipped with propulsion systems using low-thrust engines and whose postage or transfer of orbit is achieved by a number of significant orbit revolutions. These low power motors are, for example, motors whose propellant ionization is carried out in a low thrust electric manner typically using a grid ion or Hall effect nozzle technology. These engines are also known in the state of the art, under the name of electric motors. Engines whose ionization of the propellant is carried out chemically are known in the state of the art as chemical engines; they are generally intended to deliver a strong thrust but can also be used to deliver a low or medium thrust. These low thrust engines make it possible to limit the mass of fuel required to perform the satellite orbit transfer operation. However, since these motors are of low power, they have the disadvantage of an elongation of one to two orders of magnitude of the time of placement or transfer of orbit relative to the use of high thrust chemical engines. The nominal duration of low thrust orbital transfer may indeed vary from a few weeks to a few months. Due to this low power and the lengthening of the transfer or positioning time, the control methods, which determine the engine thrust law (direction and amplitude as a function of time), used for chemical engines with high thrust, are not applicable for low thrust engines. One method of controlling electric motors is described in the publication "Boeing Low-Thrust Geosynchronous Transfer Mission Experience", for the orbital transfer of an elliptical injection orbit delivered by a launch vehicle to a geostationary target orbit. It consists in a first phase to apply a continuous thrust according to the instantaneous velocity vector of the satellite until it reaches an elliptical orbit of the same period as that of the target orbit. A second phase is devoted to the transformation of this elliptical orbit into a circular orbit using a thrust orientation law perpendicular to the apogee-perigee line in the plane of the orbit. This method has some disadvantages: - it is suboptimal insofar as the transfer time is too long and the consumption of electric fuel (Xenon, Argon, ...) too important; it does not make it possible to reach the circular target orbit such as the operational orbit with sufficient precision; it is limited to a transfer of GTO-GEO type, that is to say to a transfer from an elliptical orbit to a circular orbit of 24 hours. In addition, it is also possible to aim for a transfer from a non-elliptical orbit to a non-circular orbit or more generally a transfer whatever the orbit of departure and arrival of the satellite. These disadvantages are overcome by the method of positioning described in the patent application FR 2998875. It can be carried out on board the satellite (in particular having requirements for memory and calculation resources compatible with the performance of a satellite) . This method makes it possible to determine the optimal control law regardless of the starting or arrival orbit of the satellite, while minimizing the travel time or the fuel consumption during the satellite positioning or transfer of the orbit. The resources, in terms of the amount of available memory and computing power needed to operate the process, are small compared to the computing resources of the current satellites. The control method is robust to mission interruptions, such as interruption of maintenance control, breakdowns, etc. The control method is able to automatically correct the optimal control law in a closed loop according to the deviation to the nominal trajectory, with simple calculations and without re-programming from the ground. Finally, this solution allows the realization of an autonomous transfer of orbit and is adapted to the use of electric motors.
    [0002] But this method which is based in particular on the knowledge on board the satellite and in real time, the position of the satellite, requires that it is equipped with a GNSS receiver type. Such a receiver is difficult to design because the acquisition of information is performed on the secondary lobes of the receiver antenna with a low SNR. And such a receiver is not suitable for orbits or portions of orbit whose altitude is higher than that of the constellation of GNSS satellites which is about 20 000 km. Moreover, this method, which is particularly well suited to the positioning of a satellite in self-rotation around the thrust vector, poses a problem of implementation when the satellite is not in this configuration. The object of the invention is to overcome these disadvantages. The guiding method according to the invention is based on a unique representation from the theory of optimal control (no paradigm shift). More specifically, the subject of the invention is a guiding method for setting a low-thrust satellite in operation equipped with communication means with a ground station, characterized in that it comprises the following steps performed during a current cycle. predefined: A) ground determining for a predetermined cycle, a satellite thrust vector orientation law, and a history of satellite state and satellite state variables for the transfer of a departure orbit. to a predetermined target orbit using optimal control theory, B) from the satellite thrust vector orientation law, and the history of satellite state and satellite state variables, determine for the said cycle period, a law 30 of evolution of the rotation of the satellite around the vector pushed into an inertial reference frame, C) to represent in a predetermined format the evolution of ion of the state and deputy state variables to obtain first parameters, D) to represent in a predetermined format a law of evolution of the rotation to obtain second parameters, E) concatenate the first and second parameters to obtain a first parameter, satellite guide plane; F) download the satellite guide plane on board the satellite; G) during the current cycle, iterate the following substeps according to a predefined period that is shorter than the duration of the guidance cycle: - gl) reconstruct on board the satellite, a satellite guidance set; - g2) executing on board the satellite the guidance instruction by applying a closed control loop, H) measuring the actual orbital trajectory of the satellite on the ground, I) repeating the steps A) to H) periodically from cycle to cycle, with the trajectory measured at the end of the previous cycle as starting orbit of the next cycle, until reaching the target orbit.
    [0003] This approach is generic in that it can be applied to any type of low thrust orbital transfer. The quality of the edge guidance control obtained by the proposed approach is better than that obtained by the state-of-the-art solution (reduction of skew and noise by construction): Open loop on board is reclosed by the ground over a periodic time horizon using standard ground means for measuring and filtering the orbit. The method according to the invention makes it possible to simply reconstruct on board the 3-axis satellite guidance attitude (for example represented by a standard unit quaternion), and not only the orientation law of the thrust vector. In addition, this generalization of the satellite control is done without additional approximation (no curve adjustment: on board, the process of reconstruction of the attitude does not introduce error). The method according to the invention proposes an efficient and robust method that can even withstand unscheduled interruptions in the mission, in order to implement on board a satellite the optimum control law minimizing the travel time or the fuel consumption for a period of time. given nominal path. It is effective in the sense that a registration is made by the ground at the beginning of each cycle by considering the measured orbit, as well as a terminating trajectory re-optimization which induces a performance of the implemented law close to that of the theoretical law. The cost of implementation in terms of memory and computation is low compared to the computing resources of the current satellites. In addition, no navigation means (receiver and GNSS antenna) is required on board, which reduces the complexity and cost of development of the satellite. Other characteristics and advantages of the invention will become apparent on reading the detailed description which follows, given by way of nonlimiting example and with reference to the appended drawings, in which: FIG. 1 schematically represents a flowchart of the various steps of the invention. According to the invention, FIG. 2 illustrates the differences in generating the guide plane between a method of the prior art (FIG. 2a) and that according to the invention (FIG. 2b). From one figure to another, the same elements are identified by the same references. The method according to the invention presupposes that the guide path is planned on the ground before the beginning of the low thrust transfer. The first guidance plane is then downloaded onboard the satellite for application on a limited guidance horizon (eg 7 days). During the current cycle of the orbital transfer, the method makes it possible to simply calculate the on-board guidance set and execute it in an open loop. The resetting of the orbit is done on the ground. Then a new ground planning of the guidance trajectory is performed from the measured orbit to the target orbit. The new guidance plan is then uploaded to the satellite. The process stops after running the last guidance cycle. The guiding method according to the invention is thus based on an iterative loop of edge-ground guidance (for a current cycle) which is summarized by the graph of FIG. 1 in which the steps performed on the ground are indicated in straight characters, those carried on board being indicated in italics. It comprises the following steps: A) Determine on the ground for a predetermined guidance cycle, a vector orientation law of the thrust vector of the satellite, and a history (that is to say the time evolution) of the state variables and satellite adjoint state variables, for the transfer of a starting orbit to a predetermined target orbit using the optimal control theory. It is recalled that a state vector makes it possible to characterize a dynamic system in vector form by using state variables. The state variables at a given time are quantities that completely define the state of the dynamic system at this time. These quantities are most often of physical significance. The knowledge of the state vector at any instant t makes it possible to know the state over an interval [tt by integration with respect to the time between t and t T of the dynamics of the state vector. T is an arbitrary variable representing the horizon of prediction time. The number of state variables, denoted by the letter n, is the size of the system.
    [0004] B) From the satellite thrust vector orientation law, and the history of state variables and satellite state variables, represented in an inertial reference frame, determine the ground for the guidance cycle current, a law of evolution of the rotation of the satellite around the thrust vector.
    [0005] C) Representing on the ground according to a predetermined format the temporal evolution of the state and adjoint state variables to obtain first parameters.
    [0006] D) Represent on the ground according to a predetermined format the law of evolution of the rotation to obtain second parameters. E) Concatenate the first and second parameters on the ground to obtain a satellite guidance plan.35 F) Download this guidance plane on board the satellite. G) During the current cycle iterate the following steps according to a predefined period much shorter than the duration of the guiding cycle, for example every minute for a one-week guiding cycle: - gl) reconstruct on board the satellite a guidance of the satellite, - g2) perform on board the satellite guide set by applying in a conventional manner a closed control loop. 10 H) During the current guidance cycle, measure the actual orbital trajectory of the satellite on the ground according to a predefined period, for example every 4 hours, so as to obtain at the end of the cycle a real orbital trajectory. I) Repeat the previous steps on the next cycle, with the actual trajectory measured at the end of the current guidance cycle as the starting orbit, until reaching the target orbit. We will now detail these steps. Beforehand, a starting orbit is defined as well as a target orbit. The starting orbit is, for example, the injection orbit delivered by the launch vehicle, or an intermediate transfer orbit if the first part of the transfer is conventionally performed by a high thrust chemical engine; the target orbit is, for example, the operational orbit of the satellite mission (eg the geostationary orbit), or an orbit close to it. Likewise, the guidance cycle also referred to as the guidance horizon is determined beforehand, experimentally or by ground simulation, making a compromise between cycle time (preferably long) and fuel consumption (preferably low). Step A) to determine: the satellite thrust vector orientation law, a satellite state variable history, and a satellite satellite state variable history, can use different models of satellite the spatial evolution of the satellite, as described in the patent application FR 2998875.
    [0007] A first model uses a Cartesian representation. A Cartesian representation is a representation in position and speed. This first model uses the following equations: r F perturbatil: this dt2 -113 + 1- m nor dm = go X Ism In these equations the different variables represent the following elements: r radius vector of the satellite with respect to the center Earth in meters, governor of the cosines of the thrust, F thrust of the engine (F 0) in Newtons, sP specific impulse of the engine in seconds, m mass of the satellite in kilograms, g Gravitational constant 3,986,005 E + 14 m3 / s2 for the Earth, go Standardized ground acceleration 9.80665 m / s2, covering all interfering forces affecting the trajectory of the satellite in Newtons. Disruptive forces acting on the satellite are secondary to changes in the satellite's trajectory. At first, their actions are neglected and treated as disturbances by closed-loop control. Noting the speed E12. the equations of satellite velocity dynamics can be written in the form of a system of nonlinear differential equations of the first order: dt r F = -u _ 4- - rr 3 dm = dt go x Note neither the system state vector that allows to have a Cartesian representation of the dynamics of the satellite.
    [0008] A second model uses a Keplerian representation. In this model, the equation of the dynamics of the satellite is transformed to express the movement of the satellite in terms of Keplerian elements. These Keplerian elements are the semi-major axis a, eccentricity e, the longitude of the perigee a), the longitude of the ascending node n and the true anomaly y. This modeling offers the advantage of being directly interpretable by those skilled in the art. Indeed, it directly expresses the geometrical elements of the orbit of the satellite. In addition, five of the six parameters are first integrals of the movement, which allows a simple numerical implementation. The state vector in this coordinate set is x = [a, e, i, w, û, v, m]. A third model is the equinoctial model. This model uses coordinates whose parameters are p, ex, ey, h ,, hy and 1: p is the parameter of the conic, [ex, e31 represents the eccentricity vector and {itx, h) 1 the inclination vector. The state vector in this coordinate set is x = eg, e3 'hx, h)., Rni. Unlike the Keplerian model, the state dynamics of the equinoctial model do not present any singularity, both for equatorial orbits (i = 00) and polar orbits (1 = 900). In addition, the state dynamics is valid simultaneously for elliptical and hyperbolic orbits. The parameters of the equinoctial model are expressed from the Keplerian parameters by the following equations: p = a-e2f in meters = e X co s (co n) e = e X sin (cd n) Y = tan cos (SI) 2 hy = tan s in (n) 2 = v in radian In these equations, the different elements represent: a the semi-major axis in meter the eccentricity t the inclination in radian the longitude of the ascending node in radian the argument of the perigee radian the true anomaly in radian Using the equinoctial model the equations of the dynamics of the satellite are the following equations: dp pa 1 - 2 --S dt uZ deX ae .. ZXsin1) XQ + AXS F x cos (1 ) In this equation, the different parameters without units are defined by: Z = e xcos (l) -1-e. sin (1) WA = (1 + Z) where s (1) B = ey + (1 + Z) sin (1) F = h 'sin (i) - hy cos (I) X = 1 + h + hy2 and Q, s and w are the radial, tangential and normal components of the acceleration delivered by the motor and / or the disturbing forces. Since the dynamics of the satellite are evolving slowly due to the low thrust of the engines, it is interesting to look at the dynamics in the average state parameters on an orbit instead of dealing with the instantaneous state parameters as in the previous equations. The averaging operation is carried out with the following formula: 15 = 27r 1 T = dt where f is the satellite dynamic function of the state x and the control and T the period of the orbit. The averaging provides a smoother representation of the orbit parameters that are more easily represented by polynomials. With a model of the satellite dynamics selected, the reference trajectory is now determined, based on the starting orbit, the target orbit and the satellite characteristics (total mass, total thrust, and specific pulse of the satellite). electric propulsion engines used in transfer). This determination is carried out using the theory of optimal control by applying the principle of the maximum, on the chosen model. This application of the principle of the maximum makes it possible to calculate the optimal reference trajectory on the current guidance cycle according to the criterion of optimality selected: typically trajectory in minimum time or trajectory with a fixed duration and in minimum fuel consumption. This step makes it possible to obtain an optimal trajectory according to a time t, whose representative parameters are x (t) and..., F (t). xref (t) is the state vector of the satellite dynamics (for example (t) = ip (ï), e, (f), e (t), h, (t), hy (t), (t) )]) and, 2 ,,, f (0 are the Lagrange multipliers associated with the satellite adjoint state vector in application of the minimum principle, (for example = X, (t))) "(01). then determines the parameters of the control law of the motors of the satellite, associated with the optimal trajectory determined above.This determination is made by solving the equation on the control resulting from the theory of optimal control from the state x (t) Lagrange multipliers ..i (t) The control maximizes the Hamiltonian H of the problem at all times.The parameters of the control law of the motors include: - the law of orientation of the thrust of the engine , the maximization of the Hamiltonian H with respect to the thrust orientation obtained by solving the following equation: ## EQU1 ## 'ignition of the motor S, obtained by solving the following equation: max if (6 = 0), 11 (6 = 1)) a represents the boolean determining whether the motor is on (6 = 1) or off (6 = 0) ).
    [0009] This step thus makes it possible to obtain the orientation law of the thrust vector of the engine, as well as the ignition law of this thrust. It may be noted that the choice of state variables and the use of filtering or averaging techniques has a direct impact on the ease of representation and parameterization of these data on board the satellite (see step C). In fact, it is necessary to be able to download and store on board the satellite the evolution of the satellite state and state variables, discretized in time at the time step of the computer over the duration of the guidance cycle (a week for example) , which represents a download and a very important and expensive memory storage for a satellite. Classically, these tables of data are replaced by a representation of associated parameters; it is then sufficient to download and store on board the satellite only these parameters (see step F). B) From this satellite thrust vector orientation law and satellite state and state vector state histories, represented in an inertial reference frame, the optimal law of evolution of the satellite is now determined. the rotation (= evolution of the steering angle) of the satellite around the thrust vector. In other words, it is a question of optimizing the nominal 3-axis satellite attitude and the control law of the solar generators under operational constraints (kinematics, sensors, thermal, power) on the current guidance cycle. Optimization uses a conventional non-linear constrained optimization method, the objective function to be minimized being the average solar aspect angle over the time horizon considered. C) To reduce the amount of data to be downloaded and stored on board the satellite, the data tables from step A) are replaced by a representation of associated parameters. Typically, the data of the state variables and those of the adjoint state variables are replaced by a time-dependent polynomial representation of these variables; it is then sufficient to store only the 25 polynomial coefficients (a few tens of values depending on the order of the polynomial), also called first parameters. The parameter Aex averaged has for example a polynomial representation of the form) tex (t) = 1532774. t66 + 18727.26. t5- 28021.34. t4_ 3133,043. t3 30 + 792,0076. 189,1362. t + 18.59838 with a correlation coefficient extremely close to 1: R2 = 0.9999190. This polynomial is therefore a very good approximation of λ, x over the entire duration of the path. Seven coefficients are enough to represent it.
    [0010] In the case of a classical representation without prior averaging, a single polynomial does not make it possible to represent p correctly over the entire trajectory because the solution of the problem is oscillating; in practice, it will be necessary to break the trajectory into small ends and use a model based on a polynomial by oscillation. On the other hand, the search for the minimum distance of the parameter p current with respect to the nominal parameter p is complicated by the risk of seeing several local minima and thus several solutions.
    [0011] According to an alternative, this step C) can be carried out just after step A), before step B). D) As before, in order to reduce the volume of data to be downloaded and stored on board the satellite, the data of the law of rotation evolution derived from B) are replaced by: a representation in the form of a table d sampling, a polynomial representation as shown in the previous example, or a representation in another format compatible with the desired performance. The representation format chosen for the data of this law of evolution of the rotation is independent of that which has been chosen for the representation of the data of the state variables of the reference trajectory, and of the adjoint state variables. The data of this law of evolution of the rotation are for example replaced by the parameters obtained by the following representation: a representation in the form of a sampling table with a weak temporal resolution, associated with a method of interpolation between the corresponding attitudes at two successive points in the table. For example, a time step corresponding to about ten points per orbital revolution is sufficient, which makes it possible to limit the amount of type 2 data in the guidance plane. These parameters are referred to as second parameters.
    [0012] It may be noted that selecting a representation of the satellite 3-axis attitude directly in the form of Euler angles or in the form of unitary quaternions (forms conventionally used by the onboard SCAO subsystem), has the following drawbacks: the dynamics of the satellite attitude being at a higher frequency than the dynamics of the satellite state, the quality of the approximation or the interpolation used on board is mediocre and introduces errors and biases which are avoided by choosing the other recommended representations.
    [0013] E) The first and second parameters are then concatenated. These concatenated parameters define the guidance plane for the current cycle. F) This is then downloaded from the ground station to the satellite and stored on board the satellite. Conventional network means of ground stations equipped with antennas make it possible to establish the upstream link (TC) and the downlink (TM) when the satellite is in visibility of one of the stations of the network used for the posting.
    [0014] G) Then, in a conventional manner, a substep gl) makes it possible to reconstruct the guidance setpoint on board the satellite by using the parameters of the guidance plane that have been downloaded; this step is therefore performed in open loop, without correction or catch-up during the current cycle. This guidance set includes a thrust vector orientation setpoint and a 3-axis attitude setpoint of the satellite. Then (substep g2) the guidance setpoint is executed by an attitude control subsystem and an orbit control subsystem on the satellite, including a closed-loop control mechanism around the setpoint. guide.
    [0015] These steps of reconstruction and execution of the guidance set are performed periodically during the cycle, in a predefined period (for example every minute) much less than the cycle time. This causes the satellite to move during this cycle. This displacement of the satellite is in addition impacted by various external disturbances (example solar radiation pressure, harmonic of the terrestrial potential, lunisolar attraction, ...) or internal (for example the errors of realization on the thrust of the engine) which are treated by the closed-loop control of substep g2.
    [0016] H) In parallel with the progress of these steps, that is to say during the current cycle, the actual orbital trajectory of the satellite is measured on the ground regularly, according to a predefined period, for example every 4 hours. The estimated orbit is obtained by filtering measurements acquired on the ground (using a network of ground stations) with an orbit propagation model allowing to integrate the weak thrust as well as other models of natural disturbance forces (such as the effect of the flattening of the Earth) on a given orbital arc. The filter can be sequential (least squares) or recursive (Kalman filter).
    [0017] I) When this current cycle is completed, the preceding steps are repeated periodically from cycle to cycle, taking as starting orbit the real orbit measured on the ground at the end of the preceding cycle until the completion of the orbital transfer. that is to say until the arrival on the orbital target with the desired accuracy for the mission of posting.
    [0018] The solution according to the invention makes it possible to split the global trajectory and attitude generation problem by cutting into two simpler sub-problems (trajectory generation and attitude generation) treated sequentially, using an indirect approach for the trajectory generation and a nonlinear optimization approach for attitude generation. Resolving the problem globally in a single step, for example by a non-linear constrained optimization method, leads to a large number of decision variables, which causes convergence problems and too long response times. Solving the problem by splitting into two simpler sub-problems treated sequentially but using a non-indirect approach for trajectory generation does not allow to benefit from the representation related to the theory of optimal control using the satellite's adjoint state; the other approaches are specific and / or suboptimal and / or introduce errors in the onboard representation. Similarly, solving the problem by splitting into two simpler sub-problems treated sequentially but using an approach other than nonlinear optimization for attitude generation provides a feasible solution in terms of satisfying the constraints of attitude, but a non-optimized solution. The process according to the invention differs from the patent application FR 2998875 (or '875) in the following points. Sharing of edge-to-ground tasks: The '875' posting process assumes that a 'Status Determination Module (eg with a GPS)' is available on board which returns the measured state in time. real on board.
    [0019] The method according to the invention makes it possible to dispense with such a module, and the measurement of the actual state (acquisition and processing) is done on the ground from conventional measurements (distance, angular, Doppler) using a "module". ground state determination ". Consequently, the method according to the invention does not require any Lagrange multiplier recalibration edge mechanism, which is used for calculating the control edge. Guidance Horizon and Degree of Independence: The '875 Posting Process presupposes that the guidance trajectory is planned on the ground for the duration of the orbital transfer (typically 3 to 6 months) before the start of the transfer. low thrust. During the orbital transfer, this process can then be reduced to the closed loop reference trajectory. The '875 setting method applies to the entire orbital transfer without ground iteration, but this requires the development of an additional on-board' state determination module '. The method according to the invention replaces the development of such a module by a low frequency ground registration, and a re-optimization of the guidance trajectory taking into account the effect of accumulated errors on the orbit due to disturbances. not modeled. An essential point is that the process of re-optimization of the guidance trajectory is identical to that of the initial optimization, using the same means and ground tools (no further development). Content and quantity of information in the guidance plan to be downloaded to the satellite: As can be seen in FIG. 2, the method of setting 875 requires the provision of the polynomial representation of the law of evolution of the Lagrange multipliers (adjoint state parameters) over a time horizon equal to the duration of the orbital transfer (typically 3 to 6 months). In addition, the data of a sensitivity matrix is also necessary: law of evolution of the coefficients of the matrix over a horizon of time equal to the duration of the orbital transfer. The method according to the invention does not require the provision of the polynomial representation of the law of evolution of the Lagrange multipliers (adjoint state parameters) as well as that of the polynomial representation of the law of evolution of the state variables ( for example, the equinoctial orbital elements), only over a time horizon equal to the duration of a guidance cycle (for example 7 days). -: * Reconstruction on board of the only desired orientation of the thrust / reconstruction vector aboard the complete satellite 3-axis attitude: As can be seen in FIG. 2, the '875 positioning method makes it possible to reconstruct on board the thrust vector orientation law and optionally the law of "on" and "off" commands of the low thrust nozzles used for the orbital transfer. The method according to the invention also makes it possible to reconstruct the satellite 3-axis attitude (for example represented by a standard unit quaternion). For this purpose, the following components must be added to the guidance system: A "Operational constraints 3-axis attitude optimization module" activated on the ground, - Additional data in the guidance plan to be downloaded to edge of the satellite, making it possible to model and represent the law in angular rotation around the thrust vector. A generalization of the on-board command recalculation function from the guidance plane data to return the instantaneous satellite 3-axis attitude setpoint for execution by the attitude control system.
    [0020] The guidance method according to the invention can be implemented from a satellite comprising at least one engine and an attitude control subsystem, and which comprises means for implementing the guidance method presented here. above in relation to means of one or more ground stations.
    [0021] The guiding method can for example be implemented on a generic processor, a dedicated processor, a programmable gate network also known by the English name of FPGA for Field Programmable Gate Array. This guidance method may also be implemented from a computer program product, which computer program includes code instructions for performing the steps of the guidance method. It is recorded on a computer readable medium. The support can be electronic, magnetic, optical, electromagnetic or be an infrared type of diffusion medium. Such media are, for example, Random Access Memory RAMs, ReadOnly Memory ROMs, tapes, floppy disks or magnetic or optical disks (Compact Disk - Read Only Memory (CD-ROM), Compact Disk - Read / Write (CD-R / W) and DVD).
    [0022] Although the invention has been described in connection with particular embodiments, it is obvious that it is in no way limited thereto and that it includes all the technical equivalents of the means described and their combinations if These are within the scope of the invention.
    权利要求:
    Claims (7)
    [0001]
    REVENDICATIONS1. A method of guiding for setting a satellite equipped with communication means with a ground station, characterized in that it comprises the following steps performed during a predefined current cycle: A) determining on the ground for a predetermined cycle, a satellite thrust vector orientation law, and a history of satellite state and satellite state variables for the transfer of a starting orbit to a predetermined target orbit using optimal control theory, B ) from the satellite thrust vector orientation law, and the history of satellite state and satellite state variable variables represented in an inertial reference frame, determine the ground for said cycle period , a law of evolution of the rotation of the satellite around the pushed vector, C) to represent in a predetermined format the evolution of the state and state variables to obtain first parameters, D) representing, according to a predetermined format, a law of evolution of the rotation to obtain second parameters, E) concatenating the first and second parameters to obtain a satellite guide plane, F) downloading on board the satellite the plane for guiding the satellite, G) during the current cycle, periodically iterating the following substeps according to a predefined period shorter than the duration of the guiding cycle: - g1) reconstructing on board the satellite a satellite guidance set, - g2) executing on board the satellite the guidance instruction by applying a closed control loop, H) measuring on the ground the real orbital trajectory of the satellite, I) repeating the steps A) to H) periodically cycle cycle, with the trajectory measured in end of the previous cycle as the starting orbit of the next cycle until reaching the target orbit.
    [0002]
    2. Method according to the preceding claim, characterized in that the target orbit is an operational orbit.
    [0003]
    3. Method according to one of the preceding claims, characterized in that the starting orbit is an injection orbit.
    [0004]
    4. Method according to one of the preceding claims, characterized in that the evolution of the state and state of state variables is represented using a polynomial representation.
    [0005]
    5. Method according to one of the preceding claims, characterized in that the law of evolution of the rotation is represented in the form of a sampling table.
    [0006]
    6. Method according to the preceding claim, characterized in that the current cycle is one week.
    [0007]
    A computer program product, said computer program comprising code instructions for performing the steps of the method according to any one of claims 1 to 6, when said program is executed on a computer.
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    优先权:
    申请号 | 申请日 | 专利标题
    FR1402879A|FR3030456B1|2014-12-17|2014-12-17|GUIDING METHOD FOR THE POSITIONING OF A SATELLITE|FR1402879A| FR3030456B1|2014-12-17|2014-12-17|GUIDING METHOD FOR THE POSITIONING OF A SATELLITE|
    US14/967,080| US9798008B2|2014-12-17|2015-12-11|Method of guidance for placing a satellite on station|
    EP15199460.5A| EP3034411B1|2014-12-17|2015-12-11|Method for guiding the stationing of a satellite|
    CA2915368A| CA2915368A1|2014-12-17|2015-12-15|Method of guidance for placing a satellite on station|
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