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    • 专利摘要:
    The present invention relates to a method (50) for detecting a failure of a propulsion arm (20, 21) comprising a thruster (30, 31) and a hinge (22, 23, 24), said propulsion arm being adapted to forming a pair which is connected to a state of the propulsion arm by a torque-forming function. Said method comprises: - the calculation (51) of a fault isolation space, associated with the articulation of the propulsion arm, as a function of the gradient of the torque-forming function, - the estimate (52) of a speed of rotation of the spacecraft (10), - the estimate (53) of a kinetic moment momentum residue of the spacecraft, - the computation (54) of an articular residue, associated with the articulation of the propulsion arm, by projection of the kinetic momentum residue on the failure isolation space, - the search (55) of a joint failure as a function of the articular residue.
    公开号:FR3036102A1
    申请号:FR1554236
    申请日:2015-05-12
    公开日:2016-11-18
    发明作者:Francesco Capolupo；Alexandre Falcoz
    申请人:Centre National dEtudes Spatiales CNES；Airbus Defence and Space SAS；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD The present invention pertains to the field of attitude and orbit control of spacecraft, such as satellites. More particularly, the present invention relates to the detection and isolation of failure in an attitude and orbit control system of a spacecraft comprising an articulated arm carrying a thruster, called "propulsion arm". STATE OF THE ART In a known manner, a spacecraft, such as a satellite, comprises an attitude and orbit control system implemented in particular for effecting the positioning of said satellite in its mission orbit (for example to perform a transfer from a GTO orbit to a GEO orbit) and to control the attitude and position of said satellite in its mission orbit. Such systems include in particular thrusters, chemical and / or electrical, whose thrust directions are generally steerable. For example, in the case of an orbit control of a satellite previously placed in its mission orbit, the orientation of a thruster can be controlled so as to substantially align the thrust direction with the center of the mass of the satellite, to avoid forming a disturbing couple. According to another example, in the case where a kinetic moment storage device, such as a set of flywheels (reaction wheels, gyroscopic actuators, etc.), must be de-saturated, the orientation of a thruster can be controlled so as to deliberately misalign the thrust direction with the center of mass of the satellite, to form a torque adapted to de-saturate said kinetic moment storage device.
    [0002] Nowadays, it is envisaged to mount some thrusters on articulated arms, called "propulsion arm". A propulsion arm has a connecting end, attached to a body of the satellite, and a propulsion end, opposite the connecting end, carrying a propellant. Between the connecting end and the propulsion end, the propulsion arm comprises one or more joints, each having at least one degree of freedom in rotation. In the case of several joints, they are separated by a connecting element. The implementation of an articulated arm is advantageous for several reasons. Firstly, since the thruster can be placed at a distance from the satellite, the lever arm is increased and it is possible to form larger torques, for example to de-saturate the kinetic moment storage device. In addition, the thrust direction of the thruster relative to the center of mass of the satellite can be controlled with greater accuracy, for example to avoid forming a disturbing torque during orbit control operations. Also, it may be advantageous to operate the thrusters away from the satellite body to prevent them from interfering with other satellite equipment (solar generators, antennas, etc.). By cons, the use of such propulsion arms, with several joints, increases the number of equipment likely to fail. The use of such propulsion arms thus makes it more difficult to detect joint failures as well as their insulation (ie the identification of the hinge which has a breakdown), except to provide a dedicated sensor. at each articulation of each propulsion arm. DISCLOSURE OF THE INVENTION The present invention aims to remedy all or part of the limitations of the solutions of the prior art, in particular those set out above, by proposing a solution which makes it possible to detect and isolate failures of a propulsion arm while limiting the number of sensors required to do so, or even using only pre-existing sensors. For this purpose, and according to a first aspect, the invention relates to a method for detecting failure of an attitude and orbit control system 25 of a spacecraft, said system comprising at least one propulsion arm, said propulsion arm comprising at least one propellant and at least one articulation, said propulsion arm being adapted to form a pair which is connected to a state of the propulsion arm, said state of the propulsion arm corresponding to the angular position of said articulation and the thrust force standard of said thruster, by a torque-forming function. Said failure detection method comprises: calculating a fault isolation space, associated with the articulation of the propulsion arm, as a function of the gradient of the calculated torque formation function for a reference state of the propulsion arm for a control operation, - the estimation, during said control operation, of a rotational speed of the spacecraft, - the estimation of a kinetic momentum residual as a function of the rotational speed of said spacecraft, - the calculation of an articular residue, associated with the articulation of the propulsion arm, by projection of the kinetic momentum residue on the failure isolation space, 10 - the search for a joint failure as a function of the articular residue, a hinge failure being detected if said joint residue satisfies a predefined hinge failure detection criterion. Thus, the fault detection method primarily uses an estimate of the speed of rotation of the spacecraft on itself. The speed of rotation of the spacecraft can be estimated from measurements provided by sensors (gyroscopes, accelerometers, etc.) which are generally already present in the attitude and orbit control systems of spacecraft, in particular to perform attitude control, and which furthermore do not have to be placed on the propulsion arm. From the speed of rotation of the spacecraft, it is possible to estimate the real kinetic moment of the spacecraft, and also to estimate the momentum kinetic moment, that is to say the difference between the real kinetic moment of said spacecraft and a theoretical kinetic moment that said spacecraft should have, especially considering the pairs formed by the attitude and orbit control system of said spacecraft. In addition, in order to carry out the control operation in question, the propulsion arm is placed in a reference state, around which the variations of the angular position of the articulation are in principle of low amplitude for the duration of the test. control operation considered. Therefore, the torque-forming function of the propulsion arm, which is nonlinear in nature, can be linearized around this reference state by calculating the gradient of said torque-forming function for said reference state.
    [0003] 3036102 4 This linear model makes it possible to better understand the effects of a joint failure. Indeed, the calculated gradient comprises the partial derivative of the torque forming function with respect to the angular position of said articulation. The partial derivative of said torque-forming function with respect to the angular position of the articulation is representative of the direction of the parasitic torque formed by an error in the angular position of this articulation, and thus makes it possible to calculate a space of fault isolation in which a parasitic torque is formed mainly by a failure of said joint. Therefore, the articular residue, obtained by projection of the kinetic momentum residue on the failure isolation space, makes it possible to detect a failure of the joint. In particular embodiments, the fault detection method may further comprise one or more of the following features, taken alone or in any technically possible combination. In particular modes of implementation, the propulsion arm comprising at least two joints, said method comprises: the calculation of fault isolation spaces associated respectively with the various joints of the propulsion arm, the calculation of residues articular respectively associated with the various joints of the propulsion arm, - the search for a joint failure as a function of said joint residues, a hinge failure being detected if said joint residues satisfy a predefined criterion for detecting articulation failure . Thus, if the drive arm has a plurality of hinges, a failure isolation space is calculated for each hinge, as a function of the gradient of the torque-forming function calculated for the reference state. The calculated gradient comprises in particular the partial derivatives of the torque forming function with respect to the angular positions of the various articulations. The partial derivative of said torque forming function with respect to the angular position of a hinge is representative of the direction of the parasitic torque formed by an error on the angular position of this hinge, and therefore defines an orthogonal plane. in said direction of the parasitic torque, in which an error in the angular position of this joint produces a contrario no disturbing torque. Therefore, in this plane, only failures of other joints of the propulsion arm 5 are likely to form a parasitic torque. Thus, the partial derivatives allow to calculate, for each joint, a fault isolation space in which a parasitic torque is formed mainly by a failure of said joint. For example, in the case of three joints, the intersection of orthogonal planes with partial derivatives associated with two of said joints, which corresponds to a direction orthogonal to the plane defined by said two partial derivatives, defines a failure isolation space for the third articulation, insofar as only a failure of this third joint is likely to form a disturbing torque in this direction.
    [0004] In particular embodiments, the articulation failure detection criterion is verified, for an articulation of the propulsion arm, when the articular residue associated with said articulation is at least five times greater, in absolute value, than at the same time. articular residue associated with each other articulation of the propulsion arm.
    [0005] In particular embodiments, the failure detection method comprises the calculation of a propulsion residue, associated with the propellant, according to the norm of the kinetic momentum residue, and the search for a thruster failure. depending on the propulsion residue, a propellant failure being detected if the propulsion residue verifies a predefined propeller failure detection criterion. In such a case, the fault detection method thus makes it possible to detect and isolate joint and thruster failures, the insulation being however possible only when only one device (articulation or thruster) has a failure.
    [0006] In particular modes of implementation, the search for articulation failure is performed only if no thruster failure is detected. In particular embodiments, the propellant failure detection criterion 30 is checked when the propulsion residue is greater, in absolute value, than a predefined positive threshold value. In particular embodiments, the fault detection method comprises, when a thruster failure is detected, the estimation, as a function of the propulsion residue, of an error on the thrust force norm. induced by said thruster failure. In particular modes of implementation, the control operation is an operation of a posting phase of the spacecraft in a mission orbit.
    [0007] In particular embodiments, the fault detection method comprises, when a hinge failure is detected, the estimate, as a function of the articular residue associated with said hinge for which a failure has been detected, d an error on the angular position of said articulation induced by said articulation failure.
    [0008] According to a second aspect, the present invention relates to a computer program product comprising a set of program code instructions which, when executed by a processor, configure said processor to implement a fault detection method. according to any one of the embodiments of the invention.
    [0009] According to a third aspect, the present invention relates to a device for detecting a failure of an attitude and orbit control system of a spacecraft, comprising means for measuring a magnitude representative of a speed of travel. rotation of said spacecraft and means configured to implement a failure detection method according to any one of the embodiments of the invention. According to a fourth aspect, the present invention relates to an attitude and orbit control system of a spacecraft, such as a satellite, comprising at least one propulsion arm of the spacecraft, said propulsion arm comprising at least one thruster and at least one hinge, and a failure detection device according to any one of the embodiments of the invention. In particular embodiments, the attitude and orbit control system may further include one or more of the following features, taken alone or in any technically possible combination. In particular embodiments, the thruster is an electric thruster.
    [0010] In particular embodiments, the propulsion arm comprises at least two thrusters. In particular embodiments, the attitude and orbit control system comprises at least two propulsion arms. PRESENTATION OF THE FIGURES The invention will be better understood on reading the following description, given by way of non-limiting example, and with reference to the figures which represent: FIG. 1: a schematic representation of a satellite comprising two propulsion arm, 15 - Figure 2: a diagram representing the main steps of an example of implementation of a fault detection method, - Figure 3: curves representing the behavior of joint residues in the presence of a failure Figure 4: A diagram showing the main steps of a preferred embodiment of a fault detection method. In these figures, identical references from one figure to another designate identical or similar elements. For the sake of clarity, the elements shown are not to scale unless otherwise stated.
    [0011] DETAILED DESCRIPTION OF EMBODIMENTS The present invention relates to the detection of a fault at the level of a propulsion arm of a spacecraft 10, the propulsion arm being implemented by an attitude and orbit control system. said spacecraft. In the remainder of the description, there is placed in a nonlimiting manner 30 in the case of a satellite 10 to be put to post, for the purposes of its mission, in a GEO orbit. Nothing, however, excludes, according to other examples, to consider other types of spacecraft (space shuttle, orbital station, etc.), and / or other terrestrial orbits, for example geosynchronous orbits 3036102 8, Medium Earth Orbit (MEC), Low Earth Orbit (LEO), etc. FIG. 1 is a schematic representation of a nonlimiting embodiment of a satellite 10 comprising two propulsion arms 20, 21 fixed to a body 11 of said satellite 10. In the remainder of the description, reference is made in a nonlimiting manner to where the body 11 of the satellite 10 is substantially in the shape of a rectangular parallelepiped, that is to say it has six faces two by two parallel, and in the case where the attitude of the satellite 10 station in orbit GEO is controlled, for the purpose of the mission of said satellite 10 10, so as to be placed in a setpoint attitude, called "mission attitude", in which: - a face of the body 11 of the satellite 10, designated by "face + Z ", carrying for example an instrument with a payload of said satellite 10, is directed towards the Earth; the opposite face to the + Z face, arranged on the opposite side to the Earth, is designated "-Z" face; two opposite faces of the body 11 of the satellite 10, designated respectively by "face + Y" and "face -Y", are substantially parallel to the plane of the orbit GEO, the last two opposite faces of the body 11 of the satellite 10, not shown in the figures and designated respectively "face + X" and "face -X", are substantially orthogonal to a speed vector of the satellite 10 in orbit GEO. As illustrated in FIG. 1, the two propulsion arms 20, 21 are respectively fixed to the + Y face and the -Y face of the satellite 10. Nothing, however, excludes, according to other examples, considering other positions one and / or the other of said propulsion arms 20, 21. In the example illustrated in Figure 1, each propulsion arm 20, 21 comprises a thruster 30, 31. In the following description it is placed in the case where the thrusters 30, 31 are electric thrusters (electrothermal, electrostatic, plasma, etc.). Nothing, however, excludes, according to other examples, that one or both thrusters 30, 31 are chemical propellants (cold gas, liquid propellants, etc.). In addition, each propulsion arm 20, 21 comprises, in the nonlimiting example 3036102 9 illustrated by Figure 1, three joints 22, 23, 24, each joint having a degree of freedom in rotation about an axis of rotation. . The joints 22 and 23 are interconnected by a connection 25, while the joints 23 and 24 are interconnected by a link 26.
    [0012] In preferred embodiments, for each articulated arm 20, 21, the respective axes of rotation of adjacent joints 22, 23, 24 are not parallel for each of the two pairs of adjacent joints. Thus, in the nonlimiting example illustrated in FIG. 1, each propulsion arm 20, 21 offers three degrees of freedom for modifying the thrust direction and the point of application of the thrust force of the thruster 30, 31. The articulations 22, 23, 24 make it possible to modify the direction of thrust and the point of application of the thrust force of the thruster 30, 31 with respect to a center of mass O of the satellite 10. Therefore, each propulsion arm 20 21 is adapted to form a torque which varies as a function of the state of said propulsion arm, that is to say according to the angular positions of the joints 22, 23, 24 designated respectively by 01, 02 and 03 , and according to the standard F of the thrust force of said thruster 30, 31. The torque formed by each propulsion arm 20, 21 is connected to the state of said propulsion arm 20, 21 by a formation function. predefined pair. For the purposes of the attitude control, the satellite 10 may also include one or more actuators (not shown in the figures), such as reaction wheels, gyroscopic actuators, magneto-couplers, etc., adapted to form attitude control pairs of said satellite 10. In the following description, the following applies in a nonlimiting manner in the case where the satellite 10 comprises at least three reaction wheels with respective linearly independent axes of rotation, which can in addition, they can be used as kinetic moment storage devices. The invention is however applicable to any type of actuator 30 for forming an attitude control torque of said satellite 10. The satellite attitude and orbit control system 10 also comprises a control device (not shown in the figures) which controls the joints 22, 23, 24 and the thrusters 30,31 of the propulsion arms 20, 21, as well as the reaction wheels of the satellite 10, as a function of measurements made by various sensors. For example, the satellite 10 may comprise a stellar sensor ("startracker" in the English-language literature) for measuring the attitude of said satellite 10, one or more gyroscopes 5 for measuring the speed of rotation of said satellite 10 on itself, tachometers for measuring the respective rotational speeds of the reaction wheels of the satellite 10, etc. The attitude and orbit control system of the satellite 10 also comprises a device 40 for detecting the failure of the propulsion arms 20, 21. In the example illustrated in FIG. 1, the device 40 for detecting failure is embedded in the satellite 10. Nothing, however, excludes, according to other examples, having the device 40 for fault detection integrated in a ground station, or distributed between the satellite 10 and a ground station. The fault detection device 40 comprises, for example, one or more processors and storage means (magnetic hard disk, electronic memory, optical disk, etc.) in which a computer program product, in the form of a computer program product, is stored. a set of program code instructions to be executed to implement the various steps of a failure detection method 50, which will be described hereinafter. In a variant, the device 40 for detecting a fault comprises one or more programmable logic circuits of the FPGA, PLD, etc. type, and / or specialized integrated circuits (ASIC) adapted to implement all or part of the said process steps. fault detection. In other words, the device 40 for detecting a fault comprises means configured in a software (specific computer program product) and / or hardware (FPGA, PLD, ASIC, etc.) way to implement the various steps. method 50 of failure detection. FIG. 2 diagrammatically represents the main steps of a failure detection method 50, which are: calculation of fault isolation spaces, associated respectively with the different articulations 22, 23, 24 of each propulsion arm 20 , 21, as a function of the gradient of the torque-forming function calculated for a reference state of said propulsion arm 20, 21 for a control operation, an estimate of a rotation speed of the satellite 10 during said control operation, - estimating a kinetic momentum residual as a function of the measurement of the rotational speed of said satellite 10, - 54 calculation of articular residues, respectively associated with the different articulations 22, 23, 24 of each arm propulsion 20, 21, by projection of the kinetic momentum residue on the different fault isolation spaces, 10 - 55 search for a joint failure as a function of the articular residues, a hinge failure being detected if said hinge residues satisfy a predefined hinge failure detection criterion. In the example illustrated in FIG. 1, the propulsion arms 20, 21 are placed in a reference state for a control operation carried out during a transfer phase of the satellite 10 from a transfer orbit GTO to the GEO mission orbit of said satellite 10. The control operation carried out by the propulsion arms 20, 21 in FIG. 1 corresponds, for example, to the apogee maneuver of said satellite 10, and the reference state of said arms of FIG. Propulsion 20, 21 consists in forming pushing forces FO and F1 parallel to a Z axis orthogonal to the + Z and -Z faces of the body 11 of the satellite 10, oriented from the -Z face to the + Z face. The points of application of the thrust forces FO and F1 are, in the reference state, such that the thrust forces FO and F1 are misaligned with respect to the center of mass O but form opposing pairs which cancel each other out. In the remainder of the description, So = (80, Fo) denotes the reference state So in which: - eo corresponds to the vector (00,1, e0,2, 00,3) constituted by the angular positions of reference of the articulations 22, 23, 24, - Fo corresponds to the reference standard of the thrust forces FO 30 and F1 of the thrusters 30, 31. For each propulsion arm 20, 21, the pair TDTMA formed is given, as a function of the state of said propulsion arm, by a torque forming function B of the type: ## EQU1 ## During the control operation, the variations of the angular position of each articulation 22, 23, 24 are in principle of low amplitude, and the standard of the thrust force is, for example, constant. Therefore, the angular positions e1, e2, e3 and the thrust force standard F can be written in the following form: ei = eo, i + ec, i + Aei with i = 1, 2, 3 F = Fo + AF expressions in which: - ec, i corresponds to the variation controlled around the reference angular position 00, i, so that the desired angular position corresponds to 00, i + - Aei corresponds to an error on the angular position eh induced by a possible failure of the corresponding articulation, 15 - AF corresponds to an error on the standard F of the thrust force. Therefore, the torque forming function B of each propulsion arm 20, 21, which is nonlinear in nature, can be linearized around the reference state So by calculating the gradient of said torque forming function for said reference state So TDTMA To Tc +, 513, 513, 513 so Ae2, 513 AF so 502 503_Ae3 sF AT expression in which: - To corresponds to the pair formed for the reference state So, - Tc corresponds to added torque by the variations ec, i controlled (i = 1, 2, 3) around the reference state So, 25 -, 513 / δe1, δ13 /, 502 and δ13 /, 503 correspond to the partial derivatives of the function of torque formation relative to the angular positions respectively e1, e2, e3, - EB / F corresponds to the partial derivative of the torque forming function with respect to the standard F of the thrust force, 30 - AT corresponds to the torque disruptor formed by the errors Ael, 3036102 13 062, Dei and OF. The linear model given by the expression above makes it possible to better understand the effects of each joint failure. Indeed, the partial derivative EB / E4Di calculated for the reference state So, which corresponds to a vector, is representative of the direction of the parasitic torque formed by an error on the angular position (Di (i = 1, 2, 3), and therefore defines a plane, orthogonal to said direction of the parasitic torque, in which an error on the angular position (Di does not produce any disturbing torque a contrario Therefore, in this plane, only failures of the other 10 joints of the propulsion arm are capable of forming a parasitic torque.Thus, the partial derivatives allow to calculate, for each articulation, a fault isolation space in which a parasitic torque is formed mainly by a failure of said articulation. in the case where each propulsion arm 20, 21 has three hinges 22, 23, 24, the failure isolation space of each hinge is a vector space of dimension one, which can be d defined by a vector ei calculated according to the following expression: EB EB se; Esek so EB EB se; Esek so expression in which: 20 - i # j # ke (1, 2, 3), - CD corresponds to the vector product. Therefore, for each drive arm 20, 21, the failure isolation space ei is orthogonal to the partial derivatives EB / Esei and EB / Esek (i j # k), so that only one error AE); on the angular position (Di is likely to form a disturbing pair according to the vector e1, provided, however, that the partial derivative EB / E4Di is not itself orthogonal to the vector ei.Therefore, the vector ei makes it possible to detect and to isolate a failure of the joint associated with the angular position ei only if: ei 3036102 14 is (e) = eiT The expression above thus provides a set of conditions so that the failures of the various joints can be detected and Thus, it is possible to design the propulsion arms 20, 21 for these conditions to be verified for a given reference state, it is also possible to identify reference states, among several reference states associated with different control operations, for which these conditions are verified and for which the method 50 of failure detection according to the invention we can implement.
    [0013] The failure isolation spaces ei are calculated for each hinge 22 (angular position 81), 23 (angular position 82), 24 (angular position 83) of each propulsion arm 20, 21. In this way, in the case of two propulsion arms 20, 21 each having three hinges 22, 23, 24, a total of six failure isolation spaces, designated for example by eiwn (i = 1, 2, 3) for the propulsion arm 20 fixed on the face + Y, and by eq_yi (i = 1, 2, 3) for the propulsion arm 21 fixed on the face -Y. It should be noted that the failure isolation spaces e; [. 0,1 (i = 1, 2, 3) are different from the failure isolation spaces eq_yi (i = 1, 2, 3), so it is possible to distinguish the articulation failures on the propulsion arm 20 from those on the propulsion arm 21. It should be noted that the failure isolation spaces ei depend only on the formation function of the propulsion arm 20. couple B and reference state So considered. Therefore, said fault isolation gaps ei can be calculated a priori, and stored in the storage means 40 of the fault detection device 40, for later use. As previously indicated, the failure detection method 50 comprises a step 52 for estimating the speed of rotation of the satellite 10 during the control operation. The speed of rotation of the satellite 10 is, for example, estimated as a function of measurements provided by gyroscopes, or any other sensor adapted to provide a quantity representative of said rotation speed of said satellite 10. 513 # 0 is so 3036102 15 known, the speed of rotation of the satellite 10 on itself makes it possible to estimate the real kinetic moment HR of the satellite 10, which is for example given by the following expression: HR = Jw expression in which J corresponds to the matrix d the inertia of the satellite 10, and w corresponds to the speed of rotation of said satellite 10. It is also possible to estimate, in a conventional manner, the different pairs theoretically formed by the different actuators of the satellite 10 (propulsion arm, reaction wheels , etc.). From these theoretically formed couples, it is possible to estimate a theoretical kinetic moment HT that should have said satellite 10 in the absence of failure on the propulsion arms 20, 21. The theoretical kinetic moment HT is for example estimated according to the following non-inertial reference: HT = J (To ± Tc ± TRw - (10 (HRW HR)) in which: - TRw corresponds to the torque formed by the reaction wheels, - HRw corresponds to the moment The kinetics of the reaction wheels It is also possible to take into account, if they can be estimated, the disturbing pairs TPERT suffered by said satellite 10. Where appropriate, the theoretical kinetic moment HT that should have said satellite 10 in The absence of a failure on the propulsion arms 20, 21 is for example estimated according to the following expression in non-inertial reference: HT - (To ± Tc ± TRw TRERT W 0 (HRw HR t It can then be estimated, during step 53, a momentary residue of 25 minutes AH tick corresponding to the difference between the real kinetic moment RH of said spacecraft and the theoretical kinetic moment HT, which then corresponds to the kinetic momentum induced by the disturbing torque AT: AH = HR - HT ATdt During step 54, for each propulsion arm 20, 21, the articular residues re (i = 1, 2, 3) are calculated by projecting the kinetic momentum residue AH on each failure isolation space ei.
    [0014] In preferred embodiments, the articular residue rei is further filtered by means of a pseudo-derivative filter, which is approximately equivalent to estimating the disturbing torque AT by deriving the momentum residue AH kinetic and to project said disturbing torque AT on the failure isolation space 5 ei. If necessary, the articular residue rei is for example calculated according to the following expression: Rei = De (s) eiTAH expression in which De (s) corresponds to a pseudo-derivative filter, and s corresponds to the Laplace variable.
    [0015] The articular residues rei are calculated for each articulation 22 (angular position 01), 23 (angular position 02), 24 (angular position 03) of each propulsion arm 20, 21. Thus, in the case of two arms of propulsion 20, 21 each having three hinges 22, 23, 24, a total of six articular residues, designated for example by reir, y1 (i = 1, 2, 3) for the propulsion arm 20 attached to the + Y side , and by reiFyl (i = 1, 2, 3) for the propulsion arm 21 fixed on the face -Y. Next, the failure detection method 50 comprises a step 55 of searching for a joint failure as a function of the articular residues rei, a joint failure being detected if said joint residues satisfy a predefined fault detection criterion of 'joint. As indicated above, the isolation gaps ei are such that only an error AE); on the angular position ei is likely to form a disturbing torque according to the vector ei. In addition, an error AE); on the angular position ei in principle does not form a disturbing torque according to the fault isolation spaces ei and ek (i # j # k). Therefore, in case of error AE); on the angular position ei induced by a breakdown of the associated joint, the articular residue rei will in principle be much higher, in absolute value, to the articular residues rei and rek (i j # k). Thus, the articulation failure detection criterion can for example be considered as verified when the articular residue rei is at least five times greater, in absolute value, than each of the articular residues rei and rek (i j # k). For example, a failure of the joint associated with the angular position ei is detected if the articular residue rei satisfies the following two conditions: 3036 102 17 Ire 1> Te> Ke ma40j1,1rek) expression in which i # j # k and Te and Ke are predefined positive threshold values. As previously indicated, the failure isolation spaces e; [. 0,1 (i 5 = 1, 2, 3) are different from the failure isolation spaces eq_yi (i = 1, 2, 3). Therefore, in case of failure of a joint of the propulsion arm 20 fixed on the face + Y, then the criterion of detection of joint failure will be verified by the articular residues req ± y1 (i = 1, 2, 3), but will not be verified by the articular residues reiFyl (i = 1, 2, 3) calculated for the propulsion arm 21 fixed on the face -Y of the body 11 of the satellite 10. FIG. 3 schematically represents articular residues obtained by simulation. More particularly, the part a) of FIG. 3 represents the articular residues req ± y1 (i = 1, 2, 3) calculated for the different articulations 22, 23, 24 of the propulsion arm 20 fixed on the + Y face. Part b) of FIG. 3 represents the reiFyl (i = 1, 2, 3) articular residues calculated for the different articulations 22, 23, 24 of the propulsion arm 21 fixed on the -Y face. In the example illustrated in FIG. 3, a failure of the articulation 24 (angular position 03) of the propulsion arm 20 fixed on the + Y face of the satellite 10 occurs at a time T1. As illustrated by parts a) and b) the articular residues req ± y1 and reiFyl (i = 1, 2, 3) are all less than the threshold value Te before time T1. As illustrated by part a) of FIG. 3, from the moment T1, the articular residue re3 [+ y] increases progressively in absolute value because of the failure of the articulation 24 (angular position 03 ), and exceeds the threshold value Te. The residuals re-q ± y1 and re2 [+ y] remain below the threshold value Te, because the error 0e3 does not produce a disturbing torque in the fault isolation spaces ei [+ y] and e2 [+ y]. In addition, the articular residue re3 [+ Y] satisfies the fault detection criterion from a time T2, so that an articular failure is detected on the articulation 24 (angular position 03) of the propulsion arm. 20 fixed on the face + Y.
    [0016] As illustrated by part b), from time T1, the reiFyl (i = 1, 2, 3) joint residues all increase progressively, due to the increase in kinetic momentum residue AH and since the failure isolation spaces eiFyl (i = 1, 2, 3), calculated for the propulsion arm 21 5 fixed on the -Y side, do not exhibit any particular properties with respect to the disturbing couples formed by a propulsion arm articulation failure 20 fixed on the face + Y. As can be seen from FIG. 3, the articular residues reir, y1 and reiFyl (i = 1, 2, 3) make it possible to detect and isolate the failures of the articulations 22, 23, 24 of the arms In addition, it should be noted that the articular residues rei (i = 1, 2, 3) can also be used to estimate the error Aei induced by a failure. For example, if the articular residue rei verifies the joint failure detection criterion, then the error Aei can be estimated according to the following expression: 3.0 1-0, Iseo) FIG. 4 schematically represents the main steps of FIG. a preferred embodiment of the failure detection method 50. In addition to the steps described above with reference to FIG. 2, the detection method 50 of FIG. 4 further comprises the steps of: calculating a propulsion residue rF for each propellant 30, 31, as a function of the A kinetic moment residue standard AH, - 57 search for a thruster failure as a function of the propulsion residue rF, a thruster failure being detected if the propulsion residue rF verifies a predefined thruster failure detection criterion. During step 56, rF propulsion residues associated with the propellants 30, 31 are calculated. In preferred embodiments, the propulsion residue rF is filtered by means of a derivation filter, and is by Example calculated according to the following expression: ## EQU1 ## in which DF (s) corresponds to a pseudo-derivative filter, and s corresponds to the Laplace variable.
    [0017] In practice, a propellant failure 30, 31 will generally result in a significant momentum residue AH. In addition, a propellant failure 30, 31 acts only on the norm of kinetic momentum residue AH and does not act on the direction of said kinetic momentum residue AH. Therefore, the thruster failure detection criterion is for example considered to be satisfied when the propulsion residue rF is greater, in absolute value, than a predefined positive threshold value TF. It should be noted that, in the case illustrated by FIG. 1, in which the satellite 10 comprises two propulsion arms 20, 21 each comprising a propellant 30, 31, the kinetic moment residue AH standard is not sufficient to identify the propellant. 30, 31 which is down. On the other hand, the sign of the kinetic momentum residue AH makes it possible to identify the thruster 30, 31 that has failed. In addition, it should be noted that the propulsion residue rF can also be used to estimate the AF error induced by a failure. For example, if the propulsion residue rF satisfies the propulsion failure detection criterion, then the error AF can be estimated according to the following expression: AF = rF EB sF n / a In particular modes of implementation, and as shown in FIG. 4, the search for articulation failure is performed only if no thruster failure is detected. Indeed, as shown in FIG. 4, when a thruster failure is detected (reference 570 in FIG. 4), the execution of the failure detection method 50 stops. On the other hand, if no thruster failure is detected (reference 571 in FIG. 4) then the execution of the failure detection method 50 continues with the step 54 of calculation of the articular residues re (i = 1, 2, 3). Thus, it is possible to isolate the propellant failures 30, 31 and the failures of joints 22, 23, 24. Nothing, however, excludes, according to other examples, to continue the execution of the method 50 of detection of failure when a thruster failure is detected. For example, it is also possible to estimate the error AF induced by this failure and to take into account this error AF for the calculation of the articular residues and the search for articulation failure.
    [0018] More generally, it should be noted that the embodiments and embodiments considered above have been described by way of non-limiting examples, and that other variants are therefore possible. In particular, the invention has been described by considering a satellite 10 comprising two propulsion arms 20, 21. However, the invention is applicable for any satellite comprising at least one propulsion arm. In addition, the invention has been described by considering the case of a control operation performed during a transfer phase, in this case an apogee maneuver. However, there is nothing to preclude considering other types of control operation, including during a station keeping phase of the mission orbit satellite 10, since the IS (80, i) coefficients are not zero for the reference state considered. In addition, the invention has been described by considering a propulsion arm having three joints. However, the invention is applicable since the propulsion arm comprises at least one joint. In the case where said propulsion arm comprises a single articulation, the failure isolation space is for example defined by the partial derivative of the torque forming function with respect to the angular position of this articulation.
    [0019] In the case where said propulsion arm comprises two joints, the failure isolation spaces correspond for example to the orthogonal planes to the partial derivatives δB / Esei (i = 1, 2). In another example, the failure isolation space ei is a vector corresponding to the projection of the partial derivative δB / Esei in the plane orthogonal to the partial derivative EB / Esek 25 with i # k. In addition, the invention is also applicable when the propulsion arm has a number of joints greater than three. If the drive arm has four hinges, then the failure isolation spaces correspond for example to four different vectors. In such a case, the failure isolation spaces can not all be orthogonal to each other, but can nevertheless be calculated, depending on the partial derivatives, so as to maximize the insulation capacity of the joint failures. For example, it is possible to calculate the isolation gaps 3036102 21 so as to limit, for each articulation, the disturbing pairs formed in the fault isolation spaces of the other joints. Also, the invention has been described by considering a propulsion arm comprising a single propellant. Nothing, however, excludes, according to other examples, having a propulsion arm comprising at least two thrusters, for example for redundancy purposes. Propellants of the same propulsion arm can also be of the same type (electric or chemical) or of different types.
    权利要求:
    Claims (4)
    [0001]
    CLAIMS1 - Method (50) for detecting the failure of an attitude and orbit control system of a spacecraft (10), said system comprising at least one propulsion arm (20, 21), said arm of propulsion comprising at least one propellant (30, 31) and at least one hinge (22, 23, 24), said propulsion arm being adapted to form a pair which is connected to a state of the propulsion arm, said state of the propulsion arm propulsion corresponding to the angular position (01, 02, 03) of said articulation and to the thrust force standard of said thruster, by a torque-forming function, characterized in that said method comprises: - the calculation (51) a failure isolation space, associated with the articulation of the propulsion arm, as a function of the gradient of the torque-forming function calculated for a reference state of the propulsion arm for a control operation, - estimate (52), during the said control operation, of a speed of rotation of the spacecraft (10), - the estimate (53) of a kinetic momentum residual as a function of the rotational speed of said spacecraft (10), - the calculation (54) of an articular residue, associated with the articulation of the propulsion arm, by projection of the kinetic momentum residue on the failure isolation space, - the search (55) for a joint failure as a function of the articular residue, a hinge failure being detected if said hinge residue satisfies a predefined hinge failure detection criterion.
    [0002]
    2 - Method (50) according to claim 1, characterized in that, the propulsion arm (20, 21) comprising at least two joints (22, 23, 24), said method comprises: - the calculation (51) of fault isolation spaces associated respectively with the different joints of the propulsion arm, - the calculation (54) of articular residues associated respectively with the different articulations of the propulsion arm, - the search (55) of a hinge failure in function said joint residues being detected if said joint residues satisfy a predefined articulation failure detection criterion. Method (50) according to claim 2, wherein the articulation failure detection criterion is verified, for an articulation of the propulsion arm, when the articular residue associated with said articulation is at least five times greater, in absolute value, to the articular residue associated with each other articulation of the propulsion arm. Method (50) according to one of the preceding claims, comprising the calculation (56) of a propulsion residue associated with the propellant, as a function of the norm of the momentum kinetic moment, and the search (57) of a failure propulsor depending on the propulsion residue, a propellant failure being detected if the propulsion residue verifies a predefined propeller failure detection criterion. The method (50) of claim 4, wherein the hinge failure search (55) is performed only if no thruster failure is detected. Method (50) according to one of claims 4 to 5, wherein the propellant failure detection criterion is verified when the propulsion residue is greater, in absolute value, than a predefined positive threshold value. Method (50) according to one of claims 4 to 6, comprising, when a thruster failure is detected, the estimation, as a function of the propulsion residue, of an error on the standard of the thrust force induced by said thruster failure. Method (50) according to one of the preceding claims, wherein the control operation is an operation of a phase of posting of the spacecraft in a mission orbit. Method (50) according to one of the preceding claims, comprising, when a joint failure is detected, the estimation, as a function of the articular residue associated with said joint for which a failure has been detected, of an error on the angular position of said articulation induced by said articulation failure.
    [0003]
    3 - 5
    [0004]
    4 - 10 155 - 6 - 20 7 - 25 8 - 9 - 3036102 24 10 - Computer program product characterized in that it comprises a set of program code instructions which, when executed by a processor , configure said processor to implement a fault detection method (50) according to one of the preceding claims. 11 - Apparatus (40) for detecting failure of an attitude and orbit control system of a spacecraft, comprising means for measuring a magnitude representative of a rotational speed of said spacecraft, characterized in that it comprises means configured to implement a method (50) of failure detection according to one of claims 1 to 9. 12 - Attitude and orbit control system of a spacecraft (10), such as a satellite, comprising at least one propulsion arm (20, 21) of the spacecraft, said propulsion arm comprising at least one propellant (30, 31) and at least one articulation, characterized in that it comprises a device (40) for fault detection according to claim 11. 13 - System according to claim 12, wherein the thruster (30, 31) is an electric thruster. 14 - System according to one of claims 12 to 13, wherein the propulsion arm (20, 21) comprises at least two thrusters. 15 - System according to one of claims 12 to 14, comprising at least two propulsion arms (20, 21).
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    同族专利:
    公开号 | 公开日
    EP3294631A1|2018-03-21|
    WO2016181079A1|2016-11-17|
    EP3294631B1|2019-03-13|
    FR3036102B1|2017-05-26|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    EP2660154A2|2012-05-03|2013-11-06|Thales|Propulsion system for satellite orbit control and attitude control|
    WO2015028588A1|2013-08-30|2015-03-05|Thales|Method and device for electric satellite propulsion|
    US10046867B2|2015-09-18|2018-08-14|Orbital Atk, Inc.|Maneuvering system for earth orbiting satellites with electric thrusters|
    US10357377B2|2017-03-13|2019-07-23|Institute for Musculoskeletal Science and Education, Ltd.|Implant with bone contacting elements having helical and undulating planar geometries|
    CA3069541A1|2017-07-21|2019-01-24|Northrop Grumman Innovation Systems, Inc.|Spacecraft servicing devices and related assemblies, systems, and methods|
    CN111439392B|2019-09-24|2021-12-24|上海航天控制技术研究所|Spacecraft formation position cooperative control method|
    CN110589028B|2019-09-29|2021-07-06|上海航天控制技术研究所|Autonomous mode switching method for abnormal satellite attitude maneuver|
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    优先权:
    申请号 | 申请日 | 专利标题
    FR1554236A|FR3036102B1|2015-05-12|2015-05-12|METHOD AND SYSTEM FOR DETECTING FAILURE OF A PROPULSION ARM OF A SPATIAL VEHICLE, SUCH AS A SATELLITE|FR1554236A| FR3036102B1|2015-05-12|2015-05-12|METHOD AND SYSTEM FOR DETECTING FAILURE OF A PROPULSION ARM OF A SPATIAL VEHICLE, SUCH AS A SATELLITE|
    EP16727761.5A| EP3294631B1|2015-05-12|2016-05-12|Method and system for detecting a breakdown in a propulsion arm of a spacecraft such as a satellite|
    PCT/FR2016/051117| WO2016181079A1|2015-05-12|2016-05-12|Method and system for detecting a breakdown in a propulsion arm of a spacecraft such as a satellite|
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