• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
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    • 专利摘要:
    The invention proposes a system of nozzles (100) for a satellite intended to be autorotatively stabilized in a geostationary orbit, said satellite having three X, Y and Z reference axes, the Y axis representing the North / South axis and the Z axis corresponding to a pointing direction earth. The nozzle system comprises a first set of nozzles (101) configured to maintain the stationary position of the satellite, the first set comprising an even number of electrically propelled nozzles with a preset orientation, the even number being at least 4, said nozzles being oriented according to three spatial components, and having two to two different X and Y component signs.
    公开号:FR3014082A1
    申请号:FR1302782
    申请日:2013-11-29
    公开日:2015-06-05
    发明作者:Joel Amalric
    申请人:Thales SA；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD The present invention relates generally to geostationary satellite propulsion systems, and in particular to a nozzle system and a method for controlling the geostationary satellite system. orbit and attitude control for geostationary satellite. PRIOR ART To control their orbit and their attitude, the satellites use a set of actuators, and in particular a set of nozzles.
    [0002] The nozzle system may include electrically driven nozzles or chemically propelled nozzles. In known embodiments, the satellite comprises a hybrid nozzle system, including both electric propulsion nozzles and chemical propulsion nozzles. These nozzles are used separately for maintaining the satellite station. In particular, electrically propelled thrusters are used for out-of-plane orbit control (commonly referred to as "North-South control"), while chemical-propelled thrusters are used for orbit control. (called "East-West control"), and for kinetic momentum vector control maneuvers (unsaturation of inertia wheels). However, such a hybrid system has a wet mass mass ratio launched on mass useful payload and / or the operational life of the satellite is not favorable. In new generation satellites, we are moving more and more towards "all-electric" solutions for all the elements of the satellite. This "all-electric" approach can make it possible to gain enough in mass so that the same launcher can embark two satellites. This results in lower costs for launching satellites. Nozzle systems comprising only thrusters with electric propulsion have thus been proposed. The electric nozzles have a better specific impulse than the thrusters with chemical propulsion. However, these "all electric" nozzle systems require the provision in the satellite additional mennism orientation of the thrust low angular displacement (for example, type 2-axis nozzle orientation mechanism) or mechanisms of thrust orientation with high angular deflection (for example, articulated arm 2-axis, 3-axis or more). However, these long-throw mechanisms can generate increased development complexity leading to parallelism and / or orthogonality defects and pose reliability problems. This results in a very degraded operation in the event of a functional loss of an orientation mechanism. In addition, they increase the total weight of the satellite, the complexity of the onboard software, as well as the cost of the onboard equipment.
    [0003] General definition of the invention The invention improves the situation by proposing a system of nozzles for a satellite intended to be stabilized in autorotation on a geostationary orbit, the satellite comprising three reference axes X, Y and Z, the axis Y representing the North / South axis and the Z axis corresponding to a pointing direction earth. Advantageously, the system comprises a first set of nozzles configured to maintain the position of the satellite, the first set comprising an even number of electric propulsion nozzles with a preset orientation, said even number being at least equal to 4, the nozzles being oriented according to three spatial components, and having two by two signs of different X and Y components. According to a characteristic of the invention, the position of the fixed nozzles can be chosen so that the nozzles pass near the center of gravity of the satellite while maintaining a limited torque with respect to the capacity of the inertia wheels of the satellite. The position of the nozzles may furthermore be chosen to take account of the displacement of the center of gravity of the satellite during the lifetime of the satellite.
    [0004] According to another characteristic of the invention, the nozzle system may comprise a second set of nozzles comprising at least two electrically propelled nozzles, the second set of nozzles being configured to carry out at least the satellite positioning, and the nozzles the second set being oriented substantially along the same satellite axis. In one embodiment of the invention, each nozzle of the first set may form an inclination angle θ chosen with respect to the axis Y. In another embodiment of the invention, the nozzles of the first set have angles of inclination 8 substantially identical with respect to the axis Y.
    [0005] Alternatively, the nozzles of the first set may have angles of inclination 0 different from the Y axis. The first set of nozzles may comprise: a nozzle arranged on the edge delimited by the North and East faces of the nozzle; satellite box; and / or a nozzle arranged on the edge delimited by the south and east faces of the satellite box; and / or - a nozzle arranged on the north face in the vicinity of the edge delimited by the north and west faces; and / or - a nozzle arranged on the south face in the vicinity of the edge delimited by the south and east faces of the satellite box; and / or - a nozzle arranged on the east face of the satellite box, in the vicinity of the edge delimited by the south and east faces of the satellite box; and / or - a nozzle arranged on the west face of the satellite box, in the vicinity of the edge delimited by the south and west faces of the satellite box. According to another characteristic of the invention, the nozzles of the first set are non-coplanar. In one embodiment of the invention, at least one of the nozzles of the first set forms a pivot angle α with respect to the plane YZ. In particular, the nozzles of the first set may have respective pivot angles a with respect to the different YZ plane. According to one characteristic of the invention, the first set of nozzles may comprise at least one nozzle arranged in the vicinity of an outer corner of the satellite box. According to another aspect of the invention, the satellite may comprise inertia wheels while the first set of nozzles is used to carry out the control of the kinetic momentum vector in case of desaturation of the inertia wheels. The invention further provides an orbit control and attitude control method for a geostationary satellite, comprising a nozzle system according to one of the above features, the method comprising igniting the first set of nozzles independently of the each other during retention.
    [0006] Station keeping can be performed over a given number of control days, and for each control day, the method can include placing a nozzle of the first set of nozzles at a given orbital position, applying a push duration. chosen so that the net correction of the orbital elements at the end of the day is equal to a target correction vector. The method may also include activating the nozzles of the second set of nozzles in at least one of the following phases of the satellite's life cycle: repositioning of the satellite and stable orbiting at the end of the satellite's life. According to one characteristic of the invention, the method may comprise the simultaneous ignition of the nozzles of the second set of nozzles. The nozzle system according to the embodiments of the invention thus overcomes the disadvantages of hybrid nozzle systems. In particular, it does not require the transport of chemical propulsion subsystem or chemical propellant in the satellite, and combined control North / South and East / West is more efficient and more economical propellant. Furthermore, unlike "conventional all-electric" nozzle systems, the nozzle system according to the embodiments of the invention makes it possible to dispense with the thrust-orientation mechanisms conventionally provided for in the satellite. In addition, in case of loss of an electric propulsion nozzle, the post office and station keeping are always feasible. Other features and advantages of the invention will become apparent from the following description and the figures of the accompanying drawings, in which: FIG. 1 is a diagram showing a satellite in orbit; FIG. 2 is a diagram showing the nozzle system according to a first embodiment of the invention; FIG. 3 is a diagram showing the nozzle system according to a second embodiment of the invention; and FIG. 4 is a flowchart showing a control method of the nozzle system for maintaining and controlling the kinetic momentum vector, according to one embodiment of the invention. Annex A contains a set of formulas used in describing certain embodiments of the invention.
    [0007] The drawings and appendices to the description include, for the most part, elements of a certain character. They can therefore not only serve to better understand the description, but also contribute to the definition of the invention, if any. FIG. 1 represents an example of a geostationary satellite 10, comprising a platform equipped with solar panels 12 and a payload comprising transmitting and receiving antennas. For any satellite 10 in orbit 5 around the earth 11, a reference reference 7 linked to the satellite is defined. This reference mark is constituted by the direct orthonormal trihedron constituted by the X, Y and Z axes. In FIG. 1, the X axis corresponds to a direction of flight in an orbit around the Earth, the Y axis is oriented North / South, and the yaw axis Z is orthogonal to the plane formed by the X and Y axes and corresponds to a pointing direction earth. The satellite may also include hardware and logic equipment dedicated to its operation, such as inertia wheels integrated into the satellite box (actuators for attitude control), and an attitude and orbit control system. When the satellite is stabilized 3-axis, the Z axis called yaw axis is pointed towards the Earth, the Y axis said pitch axis is perpendicular to the plane of the orbit, and the X axis says of roll, is perpendicular to the Z and Y axes and in the same direction as the instantaneous linear velocity of the satellite in its orbit, the direction of the Y axis being such that the reference (X, Y, Z) is direct. FIG. 2 schematically represents a nozzle system 100 according to one embodiment of the invention. Figure 2 shows schematically the body of the satellite 20 in the form of a rectangular parallelepiped. The points of attachment of the solar panels on their axis of rotation are represented in the form of a rectangle, for the North solar panels 120 and the South solar panels 121. According to one aspect of the invention, the nozzle system 100 according to the invention comprises an even number of electrically driven thrusters that have a preset orientation prior to launching the satellite. In Figure 2, the satellite is represented as seen by an observer on the Earth's equatorial line near the sub-satellite point (point of intersection between the Earth's surface and the line that passes through the center of the earth and the satellite). In particular, the nozzle system 100 according to the invention may comprise a first set of nozzles 101, having an even number of nozzles at least equal to 4 (for example 4, 6 or 8). The nozzles of the first set 101 have a preset orientation and are generally oriented in a position close to the center of mass of the satellite. In a preferred embodiment of the invention, the nozzles of the assembly 101 are non-coplanar. The remainder of the description will be made with reference to a first set of nozzles having 4 nozzles without limitation. The 4 nozzles of the first set are noted below N1, N2, S1 and S2. The nozzles of the first set 101 comprise two pairs of nozzles on either side of the XZ plane: the first pair of nozzles (N1, N2) is generally directed towards the North (-Y axis) and the second pair of nozzles (S1 , S2) is generally directed South (axis + Y). Moreover, the nozzles of the same pair (for example N1 and N2) oht components along the Y axis of the same sign and components along the X axis of opposite signs. Each nozzle forms, in particular, an angle of inclination e chosen with respect to the axis Y. Thus, the nozzles of the first set of nozzles have, in pairs, components along the X axis of opposite signs, which allows the correction of all the orbital elements, and also to have leverage or torque capacity on the various axes of the satellite. The 4 nozzles of the first set of nozzles 101 according to the invention are in particular used for stationary maintenance, or alternatively in a combined manner for maintaining the position and control of the kinetic momentum vector. Each nozzle of the first set of nozzles can be ignited independently of the others.
    [0008] According to one aspect of the invention, the position of the nozzles 101 may be chosen so that the nozzles pass near the center of gravity of the satellite while maintaining a limited torque with respect to the capacity of the inertia wheels of the satellite. The position of the nozzles may furthermore be chosen to take account of the displacement of the center of gravity of the satellite during the lifetime of the satellite.
    [0009] Thus, the need for a deflection mechanism for the control of the kinetic vector is replaced by a ground presetting process making it possible to avoid the carriage of the spacers of the spacecraft on the satellite. Specifically, the ground presetting is performed so that there is no more clearance on board during satellite station keeping operations and for kinetic momentum vector control.
    [0010] The nozzles of the first set of nozzles 101 are advantageously preset before the launch of the satellite, and have a fixed orientation relative to the body of the satellite 20. In the event of failure of one of the nozzles of the first set of nozzles 101, it is it is possible according to the invention to use all the nozzles of the first set of nozzles (3 in the illustrated embodiment) for station keeping, or only some of them, in the event of failure of a nozzle, on the basis of another retention strategy and an assessment of the loss of efficiency. The nozzle system 100 may further comprise a second set of nozzles 102 having at least two fixed nozzles for the other phases of the satellite life cycle (for example 2, 3 or 4 fixed nozzles), in particular the setting into orbit, the insertion in final orbit, repositioning in longitude and putting into orbit cemetery at the end of operational life. The nozzles of the second set of nozzles 102 are oriented substantially along the same satellite axis, for example the yaw axis Z. The following description will be made with reference to a second set of nozzles 102 having two nozzles, denoted R1 and R2. , as a non-limiting example. According to another characteristic of the invention, all the nozzles of the second set of nozzles can be lit simultaneously.
    [0011] The nozzle system 100 according to the invention is suitable for all phases of the life of the satellite, and in particular: - the phase of posting, which corresponds to the period from the injection by the launcher to the rallying of the final position of the satellite; - the station keeping phase, which corresponds to the nominal operation phase of the satellite; - the emergency phase, if any, which corresponds to a failure and during which the altitude of the satellite can be modified; and the deactivation or desorbitation phase, during which the satellite is sent to a so-called cemetery orbit.
    [0012] For putting into orbit, the resulting thrust vector associated with the nozzles of the second set 102 (the nozzles of the second set of nozzles are advantageously lit simultaneously) is aligned with the desired direction for the thrust vector, as calculated by a device of FIG. low thrust trajectory optimization, implanted on the ground or on board. A system of three-axis attitude guidance and control of rotating solar panels is then used in this case. For insertion into final orbit, the repositioning in longitude, the setting into graveyard orbit, the resulting thrust vector associated with the nozzles of the second set 102 is aligned along the trajectory, that is to say substantially parallel to to the satellite velocity vector in the desired tegential direction (+ 1- S). A swinging maneuver ("yaw slew" in the English language) of + 1-90 degrees may be necessary to achieve an attitude in relation to the pointing in normal mode, and a rotation maneuver in the opposite direction, of - / -F. 90 degrees to return to normal mode.
    [0013] In case of failure of one of the electro-propelled nozzles of the second set of nozzles 102, the remaining nozzles are ignited. The first-order impact on thrust time is double, while the impact on propellant consumption (or equivalent on the Delta-V speed increment) is negligible. The inclination and pivot angles α of coplanar nozzles of the first set of nozzles 101 as well as the overall number of the nozzles 101 and 102 can be adjusted before launching the satellite, for example by means of a simulator so that the first set of nozzles 101 is used for retention, including orbit control and control of the kinetic momentum vector; - The second set of nozzle 102 is used for the postage and other phases of life of the satellite. After launching the satellite, the nozzles can then maintain the orientation initially set. Thus, with the nozzle system 100 according to the invention, it is not necessary to provide additional adjustment mechanism to readjust the orientation of the nozzles during flight. This results in a significant gain in mass in the satellite and a reduction in the overall cost of the satellite. For efficient operation, the electric thrust nozzles must be substantially aligned with the satellite's center of mass at predefined times during the operational lifetime (eg quarter-of-life alignment, mid-life alignment, third-party alignment) . Also, the nozzles of the system 10 according to the invention can be adjusted to be substantially aligned with the center of mass of the satellite, before the launch of the satellite. Such an arrangement of the nozzles of the first set 101 and their use for the stationary maintenance phase thus minimizes the torques during the operational lifetime of the satellite. Advantageously, the same electric thrust nozzle technology can be used for both sets of nozzles 101 and 102, but at different points of use.
    [0014] In particular, the nozzles of the second set 102 (R1 and R2) can be chosen on operating points different from the electrical power supplied to the cathode of the electro-propelled nozzles, and in a manner compatible with the power budget of the satellite. This results in a higher thrust with a lower specific impulse than for the first set of nozzles 101 than for the second set of nozzles 102.
    [0015] Alternatively, the nozzles of the first set 101 (N1, N2, S1, S2) can be chosen to have a lower thrust with a higher specific impulse than the nozzles of the second set 102. It is possible, however, without leaving of the present invention, to use nozzles for the first set 101 having a different technology than the second set of nozzles 102, for example a technology based on the use of an ejectable powder motor. The choice of these relative proportions of the specific thrust / pulse parameters for each set of nozzles 101 and 102 may have a different effect depending on the life phase of the satellite. In particular: - For an implementation in orbit, the number of installed nozzles and the electrical power supplied can provide a reasonable transfer time, typically from 2 to 6 months (depending on the customer and the chosen launch vehicle ), at the cost of additional fuel; - For insertion into final orbit, the repositioning in longitude, the positioning at the end of life, the chosen proportions make it possible to obtain an acceleration in thrust sufficient to make safe the zone used around the ring GEO without entering the window 20 East / West neighbor (considering a radial separation of +/- 40 km, and a longitudinal and latitudinal window of +/- 70 km or +/- 0.05 degrees); and - For retention, the chosen proportions can provide a sufficient specific impulse and limit the fuel consumption for a given cost in Delta-V. This is known per se, the specific pulse (generally noted lsp) is representative of the efficiency of a propulsion system. It is defined as a quotient of two quantities, one representing the thrust of a propellant, and the other representing the product of the propellant mass flow rate by the normal value of the acceleration of gravity (or flow-weight of the propellant ejected). The specific pulse indicates how long a kilogram of propellant produces thrust to move a mass of one kilogram (about 9.81 N force) into the Earth's gravitational field. The Delta-V parameter refers to the measure of change (Delta or A) satellite speed. It is expressed as the distance traveled per unit of time (meters per second) and is calculated by subtracting the speed before the change to the speed after the change, or by integrating the thrust acceleration module for the duration of the change. the process. The Delta-V can be used to estimate the amount of propellant that is required to complete a maneuver, a change of trajectory, to reach a distant destination The nozzle system 100 has the advantage of not requiring a propulsion subsystem additional chemical. Although such a subsystem is not required, the invention is compatible with the use of such a subsystem: for example the Xenon propellant could be used as a cold gas (with a very specific pulse). low) for very rare events during the life of the satellite, such as an event known as "Failure Detection Isolation and Recovery" (FDIR), which requires "Detection and Correction of In-Flight Anomaly", which requires a weak Delta-V (of an order of magnitude from one to a few m / sec).
    [0016] The nozzles N1, N2, S1 and S2 of the first set of nozzles 101 according to the first embodiment of the invention can be arranged on the north (210), south (212), east (214) and west (216) faces. of the satellite in several ways, in particular: - On the edge delimited by the north faces 210 and east 214 (nozzle N1) and / or on the edge delimited by the faces South 212 and east 214 (nozzle S2), - On the north face 210 in the vicinity of the edge delimited by the north faces 210 and west 216 (nozzle N2), and / or on the south face 212 in the vicinity of the edge delimited by the faces south 212 and east 214 (nozzle S1 ). They may also be arranged on the east face 214 in the vicinity of the edge delimited by the south faces 212 and east 216 (for the nozzle 51), and / or on the west face 216 in the vicinity of the edge delimited by the faces. South 212 and West 216 (nozzle S2). Such a configuration is substantially in the fixed plane YZ of the satellite box. In addition, +/- X components can be added to each electric thrust nozzle to provide three-axis attitude control capability. In the case of additional +/- X components, the 4 electric propulsion nozzles of the set 101 which are used for station maintenance can be arranged near the corners of the satellite box 20, in a "transverse" configuration according to a second embodiment of the invention as shown in FIG.
    [0017] The coordinate system conventionally used in orbital mechanics for motion equations is the RSW system where: -R designates the radial direction from the center of the Earth to the satellite, - S = W x R is the tangential direction close to the direction of the speed of the satellite, and -W denotes the direction of the orbital moment vector (out of the plane).
    [0018] The stabilization of a satellite 10 along three axes (the satellite is then called "stabilized 3-axis") consists in maintaining the reference (X, Y, Z) linked to the satellite in the nearest neighborhood of the reference mark (R, W , S), choosing X = + S, Y = -W, and Z = -R. In operational orbit (GEO), the satellite is in a so-called normal mode (3-axis stabilized) so that the thrust direction of the fixed-body electro-propelled nozzles is fixed in the local orbital coordinate system. The axis direction of the electro-propelled nozzles is given by its azimuth "Azzxy" and its Elevation "Ely" in the satellite fixed body coordinate system, as measured from + Z to + X according to + Y. The direction of the thrust vector for electro-propelled thrusters is given by the following vector, on the one hand in the XYZ fixed body coordinate system, and on the other hand in the RSW local orbital coordinate system: Ethrust = - - = - - RSW - cos El sin Az - cos El cos Az - sin El - cos El sin Az + sin El - cos El cos Az - XYZ In the rest of the description, the following notations may be used to designate the nozzles of the first set 101: 25 NE to designate the nozzle N1; NW to designate the nozzle N2, SE to designate the nozzle S1, and SW to designate the nozzle S2. Those skilled in the art will understand that the notations (NE, NW, SE, SW) for the nozzles of the first set 101 are not limiting and are used only by convention to facilitate the description of certain embodiments of the invention.
    [0019] The configuration matrix Csk of dimensions 3 by 4 represents the direction of the thrust vector of the four electro-propelled nozzles (EP) in the local orbital coordinate system, RSW: csk = VNE:: NW NW SE SW, 1 RSW In the description below, the configuration matrix and its signature will be described in detail for three preferred embodiments of the invention. The electric thrust nozzle system 100 of FIG. 2 corresponds to a symmetrical minimum configuration in a single plane. In this configuration, the four thrusters electrically propelled have two signs of opposite components and has an angle of inclination 8 substantially identical to the north / south axis. The four nozzles do not have a radial component along the Z axis: the Ebuy elevation and the AZzx / y azimuth for each of the nozzles in this first embodiment of the invention are given by the following matrix: NE NW SE SW Elzx ly - (90 ° - 0) - (90 ° - 0) + (90 ° - 0) + (90 ° - 0) Azzxly 90 ° 270 ° 90 ° 270 ° 20 The nozzles are here listed in order chosen only by convention (NE, NW, SE and SW) and which will be used in the following description. Those skilled in the art will understand that this order is in no way limiting but is chosen so as to facilitate the description below. The configuration matrix Csk corresponding to this nozzle system is then: 0 0 0 0 - sin 0 + sin 0 - sin 0 + sin 0 - cos Es - cos 0 + cos 9 + cos 8_ RSW This results in the following signature : 15 0 0 0 0- - + - + - + sgn C sK RSW The configuration matrix gives the thrust vector orientations of the different nozzles in the satellite reference. Its signature gives the signs of the components of the thrust vector for each nozzle ("0" for zero component, "+" for positive component, "-" for negative component). FIG. 3 represents the nozzle system 100 according to a second embodiment of the invention. In this second embodiment of the invention, the nozzle assembly 101 has a regular tetrahedron configuration.
    [0020] In this second embodiment, the first set of nozzles N1, N2, S1 and S2 (respectively designated in this figure by NE, NW, SE and SW) comprises four non-coplanar electric propulsion nozzles (North / South axis) which each form an angle of inclination e with respect to the axis Y, substantially equal, and a pivot angle a with respect to the plane XY. The nozzles are pairwise opposed component signs, allowing control of all orbital elements and having torque capability. In particular, the angle of inclination 6 with respect to the Y axis can be between 40 and 45 degrees while the pivot angle with respect to the XY plane can be between 10 and 20 degrees. The Elzy elevation and the azimuth Azzxy for each of the nozzles of the first nozzle assembly 101 are given by the following matrix in accordance with this second embodiment of the invention: ## EQU1 ## ) - (90 ° - 0) + (90 ° - 0) + (90 ° - 0) ZX 90 ° + a 270 ° - 90 ° -6 270 ° + The corresponding configuration matrix Csk is then: CSK + sin 0 sin a + sin e sin 6 - sin 0 sin a - sin 0 sin a RSW - sin 0 cos o- + sin 0 cos a - sin cos a + sin 0 cos o- cos 6 - cos 0 + cos 9 + cos The corresponding signature in the RSW mark is for this configuration matrix: + + - sgn CsK = - - - - The person skilled in the art will readily understand that the first embodiment of the invention corresponds to an application of the second embodiment of the invention. realization with a = 0. FIG. 4 shows the nozzle system 100 according to a third embodiment of the invention. In this third embodiment of the invention, the nozzle system has a non-regular tetrahedron configuration. In this configuration, the first set of nozzles 101 comprises four electrically driven nozzles N1, N2, S1 and S2 (respectively designated in this figure by NE, NW, SE and SW) two by two symmetrical with respect to the Y axis. (North / South axis) which each form inclination angles 8i with respect to the Y axis, and respective pivot angles ai with respect to the XY plane. The angle α1 represents the angle of rotation by turning around Y from X to Z. In FIG. 3, the angle α1 is, for example, positive for the nozzle N1. According to this third embodiment, the four nozzles NE, NW, SE and SW of the first set of nozzles 101 have respective angles of inclination ONE, eNW, eSE esw different and angles of rotation aNE, aNW, OSE, asw respective different. The elevation Elzm and the azimuth Azzxiy for each of the nozzles of the first set of nozzles 101 are given by the following matrix in accordance with this third embodiment of the invention: NE NW SE SW 25 Elzxn, - (90 ° - ONE) - (90 ° - 0 Nw) + (90 ° - OSE) + (90 ° - Osw) AZzx iy 90 ° + o-NE 270 ° - ow 90 ° - CisE 270 ° + asw The corresponding Csk configuration matrix to this third embodiment is then: _ + sin ONE sin uNE + sin ONw sin aNw - sin OSE sin crsE - sin Osw sin crsw C SK = sin ONE COS Cr NE + sin ° NW cos aNW - sin OSE cos crsE + The resulting signature is the following: + + - sgn CsK = - RSW It will easily be understood by the person skilled in the art that the first embodiment of the invention will be understood by those skilled in the art. The invention corresponds to a particular application of the second embodiment with eNE = eNw = esE = esw = 0 and pivot angles OENE = 6Nw = crsE = 6svv = a. The satellite put into orbit is subject to various disturbing forces that can alter the characteristics of its orbit, such as atmospheric friction, caused by the residual terrestrial atmosphere at the satellite altitude, which induces a braking of the satellite, or the attractions of the Sun and the Moon, which have the effect of reducing the inclination of the orbital plane on the equator. The purpose of station keeping is to maintain the characteristics of the orbit in order to enable the satellite to properly fulfill its mission. Station keeping typically takes place 4 or 5 days a week by means of stationary maneuvers. If necessary, electric thrusters can also be operated on a daily basis (7 days a week).
    [0021] Those skilled in the art will readily understand that these frequencies relating to station keeping are given by way of non-limiting example. The following vector A-Ce is used to designate the target corrections of natural disturbances that act on the geostationary orbit corrections at the end of the day, expressed in the equinoctial orbital elements (non-singular elements for the circular and equatorial orbit): MC = ka Ae x DeY Aix Air The equinoctial orbital elements include: - the element a representing the half-great axis; the elements ex and ey representing the eccentricity vectors; and the elements ix and iy representing the inclination vectors.
    [0022] The parameters Aa, tex, Ley, Aix and Aiy designate the target corrections of the equinoctial orbital elements a, ex, ey, ix and iy, respectively, at the end of the weekly maintenance cycle.
    [0023] The AV vector designates the speed-incremental control cost of the equinoxial orbital elements independently of each other: A I7T = [AVa A Vex AVey A Vix A Viy 1 -V Aa V Aex V = -Ae Vaix VAiy 2 2 i 2 Y Where V is the speed of the geostationary orbit.
    [0024] In the retention strategy according to the invention, it is provided, for each of the 4 or 5 days of control, to ignite each nozzle among the nozzles of the first set 101 once for station keeping, at different orbital positions. , and applying different thrust times so that the net correction of the orbital elements at the end of the day is equal to the vector é. The application of the station keeping strategy according to the invention, without considering the possible errors, gives the following annual cost AVsk for the station keeping control in terms of speed increment: + AViy2 AV SK cos 0 In this expression, "cos0" denotes the average cosine loss due to the angles of inclination 0 of the four nozzles of the first set 101 used for station keeping. The set of nozzles 101 according to the embodiments of the invention makes it possible to correct the equinoctial orbital elements if the following condition relating to the angles of inclination 0 of the four nozzles N1, N2, S1 and S2 is satisfied: ijAVex2 + AVey2 Oin m & min = arctan + AViy2 The value of Omin can be 5-10 degrees. The maximum value of ijAVix2 + AViy2 can be 52 misec per control year. This value represents the cost of correcting the secular drift due to the luni-solar effect. As we know, the satellites must maintain, whatever their trajectory or their orbit, a very precise attitude, to ensure, according to their mission, the correct orientation of their antennas, their solar panels, the scientific instruments placed on board. The attitude thus designates the angular orientation of the satellite. The attitude of the satellite is generally controlled by internal actuators such as inertia wheels making it possible to apply an internal torque to the spacecraft and to cause rotation around one of its axes X, Y, Z, the X, Y, Z axes referring to the reference trihedron linked to the spacecraft. However, the spacecraft has a tendency to detach itself under the action of disturbing torques produced by the environment such as solar pressure, aerodynamic friction forces, electromagnetic couples, gravity gradient torques. It is therefore necessary to actively control the angular orientation of the spacecraft and to ensure a stability of this orientation along its three axes. The attitude control is ensured permanently by a control loop comprising sensors that measure the orientation of the spacecraft, an onboard computer that processes these measurements and establishes the commands that are executed by one or more actuators for counteract drift and maintain direction in a chosen direction. However, each time the wheels provide an internal torque, their speed increases until reaching a maximum speed called saturation speed. When the maximum speed is reached, the wheels of inertia can no longer perform the compensation of the drifts and the onboard computer then engages a wheel desaturation operation (called "wheels unloading" in English language). At high altitude or in geostationary orbit GEO (acronym for the corresponding English expression "Geosynchronous Earth Orbit"), the ignition of the electric propulsion nozzles of the assembly 101, provided in the context of the stationary maintenance according to the invention, can also be used to control the kinetic momentum vector, and thus achieve the desaturation of the inertia wheels. It also makes it possible to design an orbital maneuvering plane which guarantees to remain in the field of the inertia wheels.
    [0025] In this embodiment according to the invention, the nozzles of the assembly 101 are thus jointly used for maintaining the position and control of the kinetic momentum vector. FIG. 4 is a flowchart representing the different steps implemented for maintaining the position and controlling the kinetic momentum vector according to one embodiment of the invention.
    [0026] In step 400, the properties of the four daily maintenance maneuvers are calculated without taking into account the control of the kinetic momentum vector. This calculation can be implemented according to the equations A1 and A2 of Appendix A, in the nozzle configuration example of FIG. 3. The configuration of FIG. 3 is a symmetrical nozzle configuration arranged in a single plane ( same inclination angle for the 4 nozzles and no radial component along the Z axis). More precisely, to calculate the properties of the 4 impulse maneuvers, the equation system of Appendix Al can be solved (known as Gauss equations for equinoxial orbit elements). The system of equations can be written in the form of the A2 matrix of Annex A of dimensions 4 of 5 where V designates the speed of the geostationary orbit.
    [0027] Such a system of equations can be solved numerically using a non-linear optimization solver. The decision variables represent the amplitude of the 4 impulse maneuvers (AVNE, AVNw, AVsE, AVsw) and the position of the four impulse maneuvers (LNE, LNw, LsE, Lsw) expressed in the right ascension of the satellite, ie the heme element equinoctial orbital. Those skilled in the art will note that in the Al equations and the A2 matrix of Appendix A, the impulse maneuvers are considered. However, alternatively, spread-thrust maneuvers may be considered for this calculation, using the relation A4 of Annex A which gives the finite At -burn duration of the push interval ("burn" in English). the four pushes as a function of the velocity impulse Aiibu 'of the thrust interval, the mass m of the satellite and the thrust F of the electrically propelled machine. The method provides for calculating the duration of each maneuver from the knowledge of the Delta-V and the mass of the satellite, in the embodiment where the impulse maneuvers are considered. The impulse maneuvers designate the very brief thrusts and delivering a pulse of AV speed supplied during a very short time. The corresponding thrust occurs for a negligible time before the period of the orbit.
    [0028] In step 402, the properties of an additional thrust for the control of the kinetic momentum vector are calculated, without overlap with the four preplanned flare intervals.
    [0029] Step 402 thus introduces an additional thrust for the control of the kinetic momentum vector. Steps 400 and 402 are then iterated to take into account the effect of the introduction of additional thrust (step 402) and the change in the characteristics of the maneuvers on the correction of the orbital elements. In step 404, the properties of the 4 holding maneuvers and the additional thrust are updated (for example, by modifying the initial time and the final thrust time) in order to satisfy both objectives. Those skilled in the art will readily understand that the additional maneuver is not necessary every day of maneuvering. Such a process converges to a "feasible" solution of combined control (stationary hold and kinetic momentum vector). When compared to the retention solution alone, the combined control solution (position hold and kinetic momentum vector control) is characterized by additional thrust and additional Delta-V cost. It is also possible according to the invention to solve directly in a single step the combined problem of the two objectives (as explained in Appendix A). The total number of decision variables is then: 5 (orbital elements) + 3 (kinetic momentum vector components) = 8.
    [0030] The total number of unknowns is thus 8 (positions and delta-V of each daily maneuver). The combined problem can then be described as square, that is to say, not under-determined and not over-determined. Several solutions can be found for the retention problem.
    [0031] Such a problem can then be solved notably by using a non-linear least squares routine. If we add an additional thrust, the total number of unknowns becomes 10 (5 * 2). The problem then becomes under-determined. Among the possible solutions, those with a minimal Delta-V cost are preferred. Such a problem can then be solved by using a numerical optimization routine, the objective function to be minimized being the propellant mass consumed or equivalently the sum of the Delta-V modules. The objective function to minimize the sum of the amplitudes of the 4 thrusts is given by equation A3 of Appendix A: A VsK = A VNE ± A VNw ± A VsE + AVsw (A3) The finite duration of the interval The four pushes are calculated using the following approximation (A4), where m denotes the mass of the satellite and F the thrust of the electrically propelled machine: mAVbun '(A4). The invention thus proposes an electric propulsion nozzle system whose presetting makes it possible to efficiently perform all the functions of the life of a satellite, without it being necessary to provide an additional mechanism in the satellite. This results in significant weight gain and reduced costs. Of course, the present invention is not limited to the examples and embodiments described and shown, and is capable of numerous variants accessible to those skilled in the art. In particular, the invention has been described with reference to a system of 6 nozzles but may include a greater number of nozzles according to the needs inherent to the satellite. Different criteria can be applied for the choice of the number of nozzles in each set 101 or 102. For example, if a reduced transfer time is desired, it is possible to increase the number of nozzles of the set 102 (for example 4 ). Furthermore, in order to define the replacement strategy in the event of anomaly of a nozzle of the holding assembly 101, a greater number of nozzles greater than 4 may be chosen for the set of nozzles 101. description above has been made in relation to a set of nozzles 100 comprising the first set 101 and the second set 102, in certain embodiments of the invention, the nozzle system may comprise only the nozzles N1, N2, S1 and S2 of the first set, for example in the case of direct injection by a commercial launcher into the geostationary orbit that does not require a postage. In embodiments where the set of nozzles 100 includes both the first set 101 and the second set 102, those skilled in the art will appreciate that the resulting pushing direction of the set 102 may be on other satellite axes according to existing development constraints on the satellite. On the other hand, in the above description examples of values have been given for angles of inclination e and a. However, the invention applies to other values.5 APPENDIX A Resolving the retention problem for the configuration example of FIG. 3 (A1) - 2 V 2 cos L 'V 2 sin L' V 0 Aa 2 0 0 - V 0 0 a Ae 2 cos LNE Ae y Ai, VY 2 sin LNE V 0 - sin - cos V '+ 0 x1 [+ sin 9 AVNw + - cos cos L. V sin L' V cos LNE V sin L 'V 0 0 - 2 - 2 VV 2 cos Ls, 2 cos Ls' VV 2 sin Ls, 2 sin Ls 'VV 0 0 x [- sin 61 0' to VSE + 0 ± cos 0 + sin 0 + cos Vs, 0 x cos Ls, V sin L 'V cos Lm V sin LsE V 0 0 + 2sin0 - V + 2 sin cos Ls' - 2 sin V - 2 sin cos L 'V -2sin 0 sin L' V - cos cos The V - cos 61 sin L 'V + 2 sin 0 V + 2 sin 0 cos L' V - cos B sin L 'V -2sin0 V - 2 sin cos L sE V - 2 sin 0 sin L sE V + cos 9 cos L sE V + COS sin L sE VV + 2 sin 0 sin Ls 'V + cos 0 sin Ls' (A2) Aa to Aex Ae Ai, Aiy V + 2sin0 sin L Nw V - cos 0 cos L Nw V + cos cos Ls, VA VNE AVS VsE AV _ sw _ (A3) Lens function to minimize the sum of the amplitudes of the 4 thrusts: AVSK = AV NE ± AV NW AV sE AVsw (A4) Finite duration of the push interval ("burn" in English) of the four outbreaks: mA V At burn burn5
    权利要求:
    Claims (22)
    [0001]
    REVENDICATIONS1. A nozzle system (100) for a satellite to be autorotatively stabilized in a geostationary orbit, said satellite having three X, Y and Z reference axes, the Y axis representing the North / South axis and the corresponding Z axis at a pointing direction earth, characterized in that it comprises a first set of nozzles (101) configured to carry out the maintenance station of the satellite, the first set comprising an even number of electric propulsion nozzles, with a preset orientation, said an even number being at least equal to 4, said nozzles being oriented according to three spatial components, and having two by two signs of different components X and Y.
    [0002]
    2. nozzle system according to claim 1, characterized in that the position of the fixed nozzles is chosen so that the nozzles pass close to the center of gravity of the satellite while maintaining a limited torque with respect to the capacity of the inertia wheels of the satellite.
    [0003]
    3. nozzle system according to claim 2, characterized in that the position of the nozzles is further chosen to take into account the displacement of the center of gravity of the satellite during the life of the satellite.
    [0004]
    4. nozzle system according to one of the preceding claims, characterized in that it comprises a second set of nozzles (102) comprising at least two electric thrust nozzles, said second set of nozzles being configured to achieve at least the setting in the satellite station, the nozzles of the second set being oriented substantially along the same satellite axis.
    [0005]
    5. nozzle system (100) according to one of the preceding claims, characterized in that each nozzle of the first set (101) forms an angle of inclination 8 selected with respect to the Y axis.
    [0006]
    6. nozzle system according to claim 5, characterized in that the nozzles of the first set (101) have angles of inclination e substantially identical with respect to the axis Y.
    [0007]
    7. nozzle system according to claim 1 to 5, characterized in that the nozzles of the first set (101) have different inclination angles θ, with respect to the Y axis.
    [0008]
    8. nozzle system according to one of the preceding claims, characterized in that the first set of nozzles (101) comprises a nozzle arranged on the edge delimited by the north faces (210) and East (214) of the body of the satellite (20).
    [0009]
    9. nozzle system according to one of the preceding claims, characterized in that the first set of nozzles (101) comprises a nozzle arranged on the edge delimited by the faces South (212) and East (214) of the body of the satellite (20).
    [0010]
    10. nozzle system according to one of the preceding claims, characterized in that the first set of nozzles (101) comprises a nozzle arranged on the north face (210) in the vicinity of the edge delimited by the north faces 210 and West 216.
    [0011]
    11. nozzle system according to one of the preceding claims, characterized in that the first set of nozzles (101) comprises a nozzle arranged on the south face (212) in the vicinity of the edge delimited by the south faces (212). and East (214) of the satellite box (20).
    [0012]
    12. nozzle system according to one of the preceding claims, characterized in that the first set of nozzles (101) comprises a nozzle arranged on the east face (214) of the satellite box in the vicinity of the edge defined by the South (212) and East (214) sides of the satellite box.
    [0013]
    13. nozzle system according to one of the preceding claims, characterized in that the first set of nozzles (101) comprises a nozzle arranged on the west face of the body of the satellite (216) in the vicinity of the edge defined by the South (212) and West (216) sides of the satellite box.
    [0014]
    14. nozzle system according to one of the preceding claims, characterized in that the nozzles of the first set (101) are non-coplanar.
    [0015]
    15. Spray nozzle system according to one of the preceding claims, characterized in that at least one of the nozzles of the first set (101) forms a pivot angle a with respect to the plane YZ.
    [0016]
    16. A nozzle system according to one of the preceding claims, characterized in that the nozzles of the first set (101) have respective pivot angles a with respect to the different YZ plane. 25
    [0017]
    17. nozzle system according to one of claims 15 and 16, characterized in that the first set of nozzles (101) comprises at least one nozzle arranged in the vicinity of an outer corner of the satellite box (20).
    [0018]
    Nozzle system according to one of the preceding claims, characterized in that the satellite (10) comprises inertia wheels, and in that the first set of nozzles (101) is used to realize the vector control of the moment. kinetics in case of desaturation of the inertia wheels.
    [0019]
    19. An orbit control and attitude control method for a geostationary satellite, comprising a nozzle system according to one of claims 1 to 18, characterized in that it comprises igniting the nozzles of the first set independently of each other. others during retention. 15
    [0020]
    20. The method of claim 19, characterized in that the stationary maintenance is carried out over a number of days of control given, and in that the method comprises, for each day of control, the placement of a nozzle of the first set of nozzles (101) at a given orbital position, by applying a thrust time selected so that the net correction of the orbital elements at the end of the day is equal to a target correction vector.
    [0021]
    21. Method according to one of claims 19 and 20, characterized in that it comprises the activation of the nozzles of the second set of nozzles (102) in at least one of the following phases among the phases of life of the satellite : repositioning of the satellite, putting into stable orbit at the end of the satellite's life.
    [0022]
    22. The method of claim 21, characterized in that it comprises the simultaneous ignition of the nozzles of the second set of nozzles (102). 30 10
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    同族专利:
    公开号 | 公开日
    EP2878539A1|2015-06-03|
    US20150307214A1|2015-10-29|
    FR3014082B1|2016-01-01|
    EP2878539B1|2020-10-28|
    ES2855948T3|2021-09-24|
    US9878807B2|2018-01-30|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    WO1992009479A2|1990-11-30|1992-06-11|Aerospatiale Societe Nationale Industrielle|Method for controlling the pitch attitude of a satellite by means of solar radiation pressure, and a satellite, for implementing same|
    RU2124461C1|1997-11-12|1999-01-10|Акционерное общество открытого типа "Ракетно-космическая корпорация "Энергия" им.С.П.Королева|Method of control of space vehicle equipped with jet engines at body base line directed at angle relative to axes and thrust lines shifted relative to center of mass of vehicle; system used for realization of this method and jet engine module|
    US5020746A|1989-09-29|1991-06-04|Hughes Aircraft Company|Method for satellite station keeping|
    IT1245661B|1991-01-23|1994-10-06|Selenia Spazio Spa Ora Alenia|THREE-AXIS STABILIZED SATELLITE EQUIPPED WITH ELECTRIC PROPULSERS FOR ORBITAL MANEUVERS AND TRIM CONTROL.|
    US5595360A|1994-03-25|1997-01-21|Hughes Aircraft Company|Optimal transfer orbit trajectory using electric propulsion|
    US5810295A|1996-07-10|1998-09-22|Hughes Electronics Corporation|Method and apparatus for a satellite station keeping|
    US6053455A|1997-01-27|2000-04-25|Space Systems/Loral, Inc.|Spacecraft attitude control system using low thrust thrusters|
    US6607167B2|2001-02-01|2003-08-19|The Boeing Company|Spacecraft thermal shock suppression system|
    US8439312B2|2007-07-17|2013-05-14|The Boeing Company|System and methods for simultaneous momentum dumping and orbit control|
    FR3006671B1|2013-06-07|2015-05-29|Thales Sa|FOUR-MODULE PROPULSION SYSTEM FOR ORBIT CONTROL AND SATELLITE ATTITUDE CONTROL|RU2750349C2|2014-08-26|2021-06-28|Астроскейл Израэл Лтд.|Satellite docking system and method|
    ITUB20152728A1|2015-07-31|2017-01-31|D Orbit S R L|PROPULSION SYSTEM FOR SMALL ARTIFICIAL SATELLITES OF SMALL DIMENSIONS, INCORPORATING SATELLITES OF THE PROPULSION SYSTEM AND METHOD OF MANAGEMENT OF THAT PROPULSION SYSTEM|
    ES2596721B1|2016-01-15|2017-11-06|Antonio SÁNCHEZ TORRES|Pulsed electric nozzle to increase thrust in plasma space motors|
    US10625882B2|2017-03-06|2020-04-21|Effective Space Solutions Ltd.|Service satellite for providing in-orbit services using variable thruster control|
    CN109896050A|2019-03-20|2019-06-18|西北工业大学|A kind of automatically controlled vectored thrust electric propulsion device|
    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|
    法律状态:
    2015-10-23| PLFP| Fee payment|Year of fee payment: 3 |
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    2017-10-26| PLFP| Fee payment|Year of fee payment: 5 |
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    优先权:
    申请号 | 申请日 | 专利标题
    FR1302782A|FR3014082B1|2013-11-29|2013-11-29|TUYER SYSTEM AND METHOD FOR ORBIT AND ATTITUDE CONTROL FOR GEOSTATIONARY SATELLITE|FR1302782A| FR3014082B1|2013-11-29|2013-11-29|TUYER SYSTEM AND METHOD FOR ORBIT AND ATTITUDE CONTROL FOR GEOSTATIONARY SATELLITE|
    US14/553,801| US9878807B2|2013-11-29|2014-11-25|Thrust nozzle system and method for the orbit and attitude control of a geostationary satellite|
    ES14194874T| ES2855948T3|2013-11-29|2014-11-26|Nozzle system and geostationary satellite orbit and attitude control procedure|
    EP14194874.5A| EP2878539B1|2013-11-29|2014-11-26|System of nozzles and method for orbit and attitude control of a geostationary satellite|
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