Power supply assembly for a plasma thruster of a spacecraft

Abstract

The invention relates to a power supply assembly (1) for a plasma thruster (2) of a spacecraft (100) for controlling a keeper current (Ikeeper) supplied to a keeper electrode (22) of the plasma thruster. The power supply assembly takes into account a contribution to a return current (Ibody) from the plasma thruster, which is constituted by a leakage current (Ileak) flowing to an electrical ground (102) of the spacecraft. This avoids premature ageing or degradation of the plasma thruster.

Description

Technical field

[0001] This disclosure relates to a power supply assembly for a spacecraft plasma thruster. It also relates to a plasma propulsion system and a spacecraft that incorporate this power supply assembly.

[0002] It applies to all types of plasma thrusters, including gridded ion thrusters, whether continuous discharge and designated by GIT for "Gridded-Ion Thruster", or radiofrequency discharge and designated by RIT for "Radiofrequency-Ion Thruster", as well as Hall effect thrusters, designated by HET for "Hall Effect Thruster" in English.

[0003] It also applies to all types of spacecraft, including space probes and satellites, regardless of their orbital attitudes. Prior art

[0004] Plasma thrusters produce the thrust force that constitutes the propulsion of the spacecraft by emitting ions with a high ionic kinetic energy value. For this, they incorporate a suitable ion emission system, which generates an ion beam towards the outside of the spacecraft. But they also incorporate a device for electrical neutralization of this ion beam, in order to reduce an electrostatic charge that is carried by the spacecraft and that could degrade some of its onboard devices if this electrostatic charge became too high compared to the electrical potential of the surrounding space. An electron emitter that is used for this purpose of neutralizing the ion beam includes a neutralizer when the plasma thruster is of a grid ion type, GIT or RIT, or includes a hollow cathode of the plasma thruster when it is of the Hall effect type.

[0005] As is known, each of these electrical neutralization devices comprises a holding electrode, or "keeper" in English, which is arranged at an output of the electron emitter. This holding electrode is supplied with an electric current which is dedicated and controlled, called holding current and denoted I keeper . This holding current then constitutes a contribution to a return current which flows from the electron emitter to a common return point, or CRP for "common reference potential" in English, also called NRP for "Neutralizer Return Potential" or "Cathode Return Potential" depending on the type of plasma thruster. This return current, denoted I body , is the sum of several contributions. A first contribution is the thrust control current, which is delivered to the ion emission system and designated by I beam for a grid ion thruster, GIT or RIT, or by I d for a Hall effect thruster.A second contribution is the holding current I keeper which is delivered to the holding electrode of the electron emitter used to neutralize the ion beam. The two currents I beam or I d on the one hand, and I keeper on the other hand, are parameters that determine the operating point of the plasma thruster. Their values ​​are provided by the manufacturer of this plasma thruster. However, it may appear that additional adjustment of the operation of the plasma thruster is necessary, beyond the current values ​​provided by its manufacturer, depending on the type of mission for which the plasma thruster is intended, based on simulations or situational tests of the spacecraft that is equipped with the plasma thruster.

[0006] Some plasma thrusters require that the sustain current be generated for a short time at the time of plasma thruster ignition and then turned off. Other plasma thrusters require that the sustain current be maintained for the duration of a propulsion phase, regardless of that duration. The distinction between these two requirements comes from the magnitude of the thrust control current flowing through the neutralizer or hollow cathode, depending on the type of plasma thruster. The higher this magnitude, the more likely the plasma thruster is to operate without sustain current. Conversely, the lower the magnitude of the thrust control current flowing through the neutralizer or hollow cathode, the more likely the plasma thruster is to require sufficient sustain current.In general, it is desired that the sum of the value of the thrust control current with that of the holding current has a substantially constant result, regardless of the operating point selected for the plasma thruster. Thus, the manufacturer of the plasma thruster provides a value to be adopted for the holding current I keeper which is different for each prescribed value of the thrust control current I beam or I d . Typical numerical values ​​are for example: between 2.2 A (ampere) and 3.8 A for the thrust control current, and between 3.1 A and 4.6 A for the holding current.

[0007] The operating instructions for a plasma thruster that are provided by its manufacturer to the manufacturer of the spacecraft in which this plasma thruster is used also include a setpoint value for the return current I body that returns from the plasma thruster from the electron emitter to the common return point CRP of its electrical power supply assembly. A value for this return current that is higher than the corresponding setpoint from the manufacturer is likely to cause premature aging or degradation of the plasma thruster, and in particular of its electrical neutralization device. However, in practice, an accentuated aging of the plasma thruster is observed in certain cases. An aim of the present invention is then to further improve the precision in the control of the plasma thruster.

[0008] An article by Ghislanzoni et al. "Hall Effect Thruster Direct Drive PPUs, Experimental Investigation of the Cathode Potential Grounding Problem", E3S Web Conf. Volume 16, 23-05-2017, DOI: 10.1051 / e3sconf / 20171615002, XP093089519, describes a power supply assembly for a plasma thruster according to the preamble of claim 1. Technical problem

[0009] From this situation, a general object of the present invention is to preserve or maximize a lifetime of a plasma thruster that is used on board a spacecraft.

[0010] More specifically, the invention aims to prevent premature wear or degradation of the plasma thruster. Summary of the invention

[0011] To achieve at least one of these aims or another, a first aspect of the invention provides a novel power supply assembly for a spacecraft plasma thruster, adapted to transform electrical power that is received from a power bus of the spacecraft, into electrical voltages and currents that are supplied by this power supply assembly to the plasma thruster, so that the latter generates an ion beam during an operation of the plasma thruster. The power supply assembly comprises a common return point for collecting a return current from the plasma thruster during its operation. The plasma thruster comprises an electron emitter that is intended to neutralize the ion beam during operation, and a holding electrode that is arranged at an output of the electrons emitted by the electron emitter.The electrical currents supplied by the power supply assembly include at least one holding current which is supplied to the holding electrode and which constitutes a contribution to the return current.

[0012] According to the invention, the power supply assembly is adapted to receive or produce successive evaluations of the return current or a leakage current flowing from the common return point to an electrical ground of the spacecraft. Such evaluations of the return current or the leakage current are received by the power supply assembly when the means for obtaining these evaluations are external to the power supply assembly. However, it is preferable that such means are internal to the power supply assembly so that it is autonomous for evaluating the return current or the leakage current.

[0013] According to a characteristic of the invention, the power supply assembly is arranged to adjust in real time during operation, the holding current according to the evaluations of the return current or the leakage current, in accordance with a current setpoint value.

[0014] Generally, the current setpoint value which is used according to the invention to adjust the holding current in real time during operation, may be relative to the return current or to a sum of the holding current and the leakage current.

[0015] In general, the current setpoint value used according to the invention can also be prescribed by the manufacturer of the plasma thruster. It can vary in accordance with several operating points provided by this manufacturer for the plasma thruster. In different embodiments of the invention, this current setpoint can relate either to the return current or to a partial return current consisting of the holding current and the leakage current. In the latter case, the setpoint for the partial return current can be obtained by subtracting a thrust control current setpoint from a return current setpoint.

[0016] Advantageously, the present invention makes it possible to prevent the return current, which flows from the electron emitter of the plasma thruster to the common return point of the electrical power supply assembly of the plasma thruster, from exceeding the setpoint prescribed by the manufacturer of this plasma thruster.

[0017] The invention therefore proposes to replace the use of a holding current setpoint possibly provided by the manufacturer of the plasma thruster by an adjustment of the holding current in order to comply with the return current setpoint or the partial return current setpoint.

[0018] The electrical ground of the spacecraft may be the spacecraft chassis. Return current flows from the electron emitter to the common return point, and leakage current flows from this common return point to the electrical ground of the spacecraft. For most embodiments of the invention, the return current is a combination of the thrust control current that is supplied to the ion emission system of the plasma thruster, the sustain current, and the leakage current.

[0019] The inventor has found that the leakage current, denoted I leak in the following, results from variations in an electrical charge that the spacecraft captures depending on various parameters such as the surface area of ​​its solar panels, their orientation, the bus voltage of the solar panels, the ion flux that is emitted by the plasma thruster, the solar illumination, the distance from the Sun, etc. This leakage current therefore constitutes a contribution to the return current that flows from the electron emitter to the common return point, in addition to the contributions of the thrust control current and the holding current. The leakage current then flows through a dedicated discharge conductive system, from the common return point to the electrical ground, and is expected to be of the order of 0.08 A to 0.1 A.Thus, without the present invention, the leakage current may cause the return current to be either totally or partially over or under the setpoint prescribed by the manufacturer of the plasma engine. The present invention therefore prevents such an over or undershoot from occurring, thereby preventing premature aging of the plasma thruster and preserving its proper operation. By virtue of the invention, the holding current may be reduced or increased by the power supply assembly, depending on the sign and / or variations of the leakage current.

[0020] Preferably, the power supply assembly may be arranged to adjust the holding current in real time during operation according to a feedback loop that comprises a subtractor, a first input of the subtractor being arranged to receive the current setpoint value, and a second input of the subtractor being arranged to receive the return current evaluations or to receive a sum of a measurement result of the holding current and the leakage current evaluation. The power supply assembly is then configured to adjust the holding current according to a difference result that is output by the subtractor.

[0021] In first embodiments of the invention, the power supply assembly comprises a measuring system which is arranged to deliver successive leakage current measurement results. In this case, each leakage current measurement result constitutes an evaluation thereof. The power supply assembly may then further comprise a summer which is arranged to receive on two inputs of this summer, at each of a series of successive times during operation of the plasma thruster, the leakage current measurement result and the holding current measurement result for this time. An output of the summer is then connected to the second input of the subtractor.

[0022] In second embodiments of the invention, the power supply assembly further comprises the measuring system for delivering the successive leakage current measurement results. The power supply assembly may then further comprise a summer which is arranged to receive on three inputs of this summer, at each of a series of successive instants during operation of the plasma thruster, the leakage current measurement result and the holding current measurement result for the same instant, and also an evaluation of a thrust control current which is supplied at this instant by the power supply assembly to the ion emission system of the plasma thruster. The output of the summer is again connected to the second input of the subtractor.

[0023] For these first two embodiments, the measuring system which is used to deliver the successive leakage current measurement results may be of a type suitable for carrying out electric current measurements.

[0024] Alternatively, the power supply assembly may further comprise a discharge conductive system which is intended to electrically connect the common return point to the electrical ground of the spacecraft. In this case, the measuring system which is used to deliver the successive leakage current measurement results may be of a type adapted to carry out electrical voltage measurements, and be arranged to measure an electrical voltage which exists between terminals of at least a part of the discharge conductive system. The power supply assembly is then configured to provide each leakage current measurement result as a result of a division of the electrical voltage measured between the terminals of the part of the discharge conductive system by a resistance value of this part of the discharge conductive system.Advantageously, the power supply assembly can be autonomous in order to have the resistance value of the part of the discharge conductive system at the terminals of which the electrical voltage is measured.

[0025] In third embodiments of the invention, the power supply assembly comprises a measuring system which is arranged to deliver successive return current measurement results, each return current measurement result constituting an evaluation thereof. The power supply assembly is then arranged to transmit these successive return current measurement results to the second input of the subtractor during operation of the plasma thruster.

[0026] Generally advantageously for the invention, the power supply assembly may be configured to adjust the holding current as a function of the difference result that is output by the subtractor in accordance with a proportional-integral transfer function that is applied to that difference result.

[0027] A second aspect of the invention provides a plasma propulsion system for a spacecraft, which comprises: at least one plasma thruster, said plasma thruster comprising an electron emitter and a holding electrode which is disposed at an output of the electrons emitted by the electron emitter; and a power supply assembly which is in accordance with the first aspect of the invention.

[0028] This plasma propulsion system is arranged so that the power supply assembly provides to the holding electrode of the plasma thruster, during operation of the latter, a holding current which is adjusted in real time according to the evaluations of the return current or the leakage current, in accordance with the current setpoint value.

[0029] When the feedback loop including the subtractor is used, the holding current which is supplied by the power supply assembly to the holding electrode can be in accordance with the difference result which is output by the subtractor.

[0030] In such a plasma propulsion system according to the invention, the plasma thruster may be of one of the following types: a grid-type, continuous discharge, or radio frequency ion thruster, wherein the electron emitter comprises an electron gun that is intended to neutralize the ion beam during operation of the plasma thruster, and the holding electrode is disposed at an electron outlet of the electron gun; or a Hall effect thruster, wherein the electron emitter is a neutralizing hollow cathode of the Hall effect thruster, and the holding electrode is disposed at an outlet opening of the hollow cathode.

[0031] A third aspect of the invention provides a spacecraft that includes a plasma propulsion system according to the second aspect of the invention.

[0032] Finally, a fourth aspect of the invention provides a method for powering a plasma thruster on board a spacecraft, this method comprising the following steps which are carried out during movement of the spacecraft in space: providing a holding current to a holding electrode that is disposed at an electron output of an electron emitter of the plasma thruster; collecting successive evaluations of a return current flowing from the plasma thruster to a return common point, or a leakage current flowing from the return common point to an electrical ground of the spacecraft; collecting a current setpoint; and adjusting the holding current that is supplied to the holding electrode based on the evaluations of the return current or the leakage current, and in accordance with the current setpoint.

[0033] In some implementations of the method of the invention, the holding current that is supplied to the holding electrode may be adjusted so that the return current, or a sum of a measurement result of the holding current and the evaluation of the leakage current, corresponds to the current setpoint value.

[0034] Thus, the holding current that is supplied to the holding electrode at a time of operation of the plasma thruster is adjusted based on evaluations or measurements that have been collected for the holding current, and for the return current or leakage current.

[0035] The current setpoint can be prescribed by a plasma thruster manufacturer. It is relative to the return current or the sum of the holding current and the leakage current, also called partial return current.

[0036] Such a method may be implemented using a power supply assembly that is in accordance with the first aspect of the invention, including the improvements that have been cited. This power supply assembly is on board the spacecraft and arranged to supply electrical voltages and currents to the plasma thruster, so that the latter generates an ion beam so as to apply a desired thrust force to the spacecraft.

[0037] The method of the invention may include subtracting the estimate of the return current, or the estimate of the sum of the holding current and the leakage current, from the current setpoint, and then using a result of the subtraction to adjust the holding current.

[0038] In various implementations of the method of the invention, one of the following features may be reproduced, alone or in combination of several of them: each evaluation of the leakage current can be obtained by a measurement of this leakage current; each evaluation of the return current can be obtained by a measurement of this return current, or by calculating a sum of a measurement result of the holding current with a measurement result of the leakage current, and with an evaluation of a thrust control current which is supplied to the ion emission system of the plasma thruster; the evaluation of the thrust control current can be formed by a set value which is used for this thrust control current, in particular prescribed by the manufacturer of the plasma thruster, or can be obtained by a measurement of this thrust control current; and the holding current can be adjusted by applying a proportional-integral transfer function to the subtraction result. Brief description of the figures

[0039] The characteristics and advantages of the present invention will appear more clearly in the detailed description below of non-limiting exemplary embodiments, with reference to the appended figures among which: [ Fig. 1 ] shows, in a schematic and simplified manner, a spacecraft in which the present invention is used; [ Fig. 2 ] is a diagram which illustrates a first possible embodiment of the invention; [ Fig. 3 ] corresponds to [ Fig. 2 ] for a second possible embodiment of the invention; and [ Fig. 4 ] corresponds to [ Fig. 2 ] for a third possible embodiment of the invention. Detailed description of the invention

[0040] In these figures, all elements are represented only symbolically, and identical references indicated in different figures designate identical elements or which have identical functions.

[0041] In [ Fig. 1 ], reference 100 designates a spacecraft, regardless of the type of this craft, for example a satellite or a space probe. References 101 and 102 designate respectively the electrical power bus and the electrical ground of the spacecraft 100. In general, the electrical ground 102 is constituted by a chassis of the spacecraft 100 on which all the on-board components are fixed.

[0042] Inside the spacecraft 100, the reference 10 denotes a plasma propulsion system, commonly designated by the acronym PPS for “Plasma Propulsion Subsystem” in English. This plasma propulsion system 10, which is sometimes referred to as a subsystem with respect to the spacecraft 100, itself comprises at least one electrical power supply assembly 1, commonly designated by the acronym PPU for “Power Processing Unit” in English, and at least one plasma thruster 2. The electrical power supply assembly 1 and the plasma thruster 2 which are represented in the figure are associated with each other so that the plasma thruster 2 is supplied with electrical energy appropriately by the electrical power supply assembly 1, in order to produce the thrust force which is desired at each instant of operation of the plasma thruster.The electrical power supply assembly 1 is connected between the electrical power bus 101 and the electrical ground 102 of the spacecraft 100, while the various components of the plasma thruster 2 are supplied with electrical currents and voltages via the electrical power supply assembly 1.

[0043] For clarity of this description, the components of the plasma propulsion system 10 that are shown in [ Fig. 1] are limited to those covered by the invention. In the plasma thruster 2, the reference 21 designates its ion emission system, regardless of the type of the plasma thruster 2 among a continuous discharge grid ion thruster designated by GIT for “Gridded Ion Thruster”, a radiofrequency discharge grid ion thruster designated by RIT for “Radiofrequency Ion Thruster”, and a Hall effect thruster designated by HET for “Hall Effect Thruster”. The reference 23 designates an electron emitter whose function is to neutralize the ions which are emitted by the system 21. For a GIT or RIT grid ion thruster, the electron emitter 23 is a neutralizer which is separate from the ion emission system 21. For an HET Hall effect thruster, the electron emitter 23 is constituted by a hollow cathode associated with the ion emission system 21. In a known manner, a holding electrode 22, or “keeper” in English, is associated with the electron emitter 23.The design and shape of this holding electrode 22 depend on the type of plasma thruster 2, but its function and use in the present invention are the same regardless of the type of plasma thruster. The holding electrode 22 has the function, in particular, of ensuring operational readiness of the electron emitter 23, by keeping the plasma sufficiently hot. For this, the holding electrode 22 is arranged close to the outlet of the electrons which are generated by the emitter 23. Generally, the holding electrode 22 physically surrounds an outlet opening of the electrons which are emitted, so as to create a holding plasma at this location, both for a GIT or RIT thruster neutralizer and for a hollow cathode of an HET thruster.

[0044] Within the electrical power supply assembly 1, an interface 11 is dedicated to electrically supplying the ion emission system 21, with a current which is commonly denoted I beam when the plasma thruster 2 is of the GIT or RIT grid ion type, or denoted I d to designate a discharge current when the plasma thruster 2 is of the HET Hall effect type. In the following, this supply current of the ion emission system 21 is generically called the thrust control current, and denoted I beam / ld or I beam ;I d .

[0045] Also inside the power supply assembly 1, an interface 12, which is distinct from the interface 11, is dedicated to supplying the holding electrode 22 with another current, which is called holding current and commonly noted I keeper.

[0046] Then, during operation of the plasma propulsion system 10, a return current which is noted I body flows from the electron emitter 23 to a common return point which is commonly designated by CRP, for “Common Return Point” in English. This common return point CRP is itself electrically connected to the electrical ground 102 of the spacecraft 100 by a discharge conductive system 103, sometimes called “bleed conducting system” or “bleed resistor” in English. I leak then designates a leakage current which flows from the common return point CRP to the electrical ground 102, through the discharge conductive system 103. A measurement system 104 can be arranged at the discharge conductive system 103, to measure in real time the leakage current I leak.For example, the measuring system 104 may be a voltmeter which is connected across the discharge conductive system 103, and the instantaneous value of the leakage current I leak may then be calculated by dividing a voltage measurement result which is delivered by the voltmeter by the known ohmic value of the discharge conductive system 103. Preferably, the systems 103 and 104 may be integrated into the power supply assembly 1 to enable operation of this power supply assembly which is autonomous for measuring the leakage current I leak.

[0047] For the architecture and operation just described, the power supply interface 11 of the ion emission system 21 has a first output terminal 11a which is connected to the ion emission system 21, to deliver to the latter the thrust control current I beam / I d . The interface 11 also has a second output terminal 11b which is connected to the common return point CRP to recover from this point the current I beam / I d . Likewise, the interface 12 has a first output terminal 12a which is connected to the holding electrode 22 to deliver to the latter the holding current I keeper . It also has a second output terminal 12b which is connected to the common return point CRP to recover the current I keeper from this point. Under these conditions, the return current I body is the sum of the thrust control current I beam / I d , the holding current I keeper and the leakage current I leak. In other words: I body = I beam + I keeper + I leak for a GIT or RIT type 2 plasma thruster, or I body = I d + I keeper + I leak for a HET type 2 plasma thruster.

[0048] The values ​​of the two currents I beam / I d and I keeper to be adopted are generally prescribed by the manufacturer of the plasma thruster 2, for one or more operating points of the latter depending on the thrust force that is desired. Furthermore, the manufacturer also provides a current setpoint value I boby_max_crit for the return current I body , or a current setpoint value I warm_crit for the sum I keeper + I leak of the holding current and the leakage current. But these setpoint values ​​are provided assuming in both cases that the leakage current I leak is zero or negligible. However, the leakage current I leak depends on many parameters external to the spacecraft 100, and / or the orientation and / or the position of this spacecraft in space, etc. Because of this, the value of the leakage current I leak is difficult to predict, as it can be positive or negative.The leakage current I leak can then cause significant deviations between the actual value of the return current I body and the set value I boby_max_crit , or between the actual value of the sum I keeper + I leak and the set value I warm_crit . Such deviations are likely to cause premature aging or degradation of the plasma thruster 2.To avoid such a risk of premature aging or degradation, the invention which is the subject of the present description proposes to add a control loop to the electrical power supply assembly 1, to adapt in real time the value of the holding current I keeper in order to ensure that at each instant of operation of the plasma thruster 2 in space, that is to say whatever the instantaneous value of the leakage current I leak , the instantaneous value of the return current I body remains substantially equal to the set value I boby_max_crit which is prescribed by the manufacturer, or else in order to ensure that the sum I keeper + I leak of the holding current and the leakage current remains substantially equal to the set value I warm_crit which is prescribed alternatively by the manufacturer. A circuit which produces such control can in particular be incorporated in the electrical power supply interface 12 of the holding electrode 22.In the general part of this description, the sum I keeper + I leak of the holding current and the leakage current has been called the partial return current.

[0049] The setpoint value I boby_max_crit or I warm_crit , as the case may be, may be supplied to a subtractor 120 which is used in the embodiments described below, by suitable setpoint recovery means 111. These setpoint recovery means 111 may be means for reading and selecting setpoint values ​​which have been initially recorded on a medium (not shown) on board the spacecraft 100, before the latter is launched into space. Preferably, this recording medium and the setpoint recovery means 111 may be integrated into the power supply assembly 1, and are activated by an on-board computer of the spacecraft 100. This on-board computer is designated by the reference 110 in [ Fig. 1] and commonly referred to as OBC for “On-Board Computer” in English. The setpoint values ​​can be recorded for several operating points of the plasma thruster 2. In this case, the reading and selection means transmit the setpoint values ​​which correspond to an operating point specified by an operator of the spacecraft 100, in order to produce the desired thrust force on this spacecraft 100.

[0050] The embodiment of the invention which is illustrated by [ Fig. 2] corresponds to the case where the values ​​prescribed by the manufacturer are the value of the thrust control current I beam / I d and the current setpoint value I warm_crit for the sum I keeper + I leak of the holding current and the leakage current. It is possible that the manufacturer prescribes a setpoint value for the holding current I keeper , but this is reinterpreted according to the invention as the setpoint value I warm_crit relating to the sum I keeper + I leak . In [ Fig. 2 ], reference 12 also designates the power supply interface of the holding electrode 22 as known before the present invention and presented previously with reference to [ Fig. 1]. Reference 120 designates a subtractor, reference 121 designates a summer, and reference 122 designates a system for measuring the holding current I keeper which is actually delivered by the interface 12 to the holding electrode 22. The measuring system 122 can be arranged on the connection which electrically connects the output of the interface 12 to the holding electrode 22, but other arrangements are alternatively possible. The summer 121 receives on its two inputs on the one hand a result of the measurement of the leakage current I leak which is delivered by the system 104 and denoted M(I leak ), and on the other hand a result of the measurement of the holding current I keeper which is delivered by the system 122 and denoted M(I keeper ). The adder 121 thus provides as output an evaluation of the sum of the holding current I keeper and the leakage current I leak , noted E(I keeper + I leak ) in [ Fig. 2]. The subtractor 120 receives the setpoint value I warm_crit on its positive input, and the summation result which is provided by the adder 121 on its negative input. The difference Δ which is calculated by the subtractor 120 is transmitted to a control input of the interface 12, and is used by the latter as an adjustment increment of the holding current I keeper to be delivered. In this way, the holding current I keeper is adjusted in real time during the operation of the plasma thruster 2, so that at each of a series of instants during this operation, the sum of the instantaneous values ​​of this holding current I keeper and the leakage current I leak remains substantially equal to the setpoint value I warm_crit .

[0051] In the embodiment of the invention which is illustrated by [ Fig. 3], the current setpoint value that is provided by the manufacturer or considered to control the power supply interface 12 of the holding electrode 22 is I body_max_crit which relates to the return current I body . The two-input adder 121 is replaced by a three-input adder 121' which respectively receive the result M(I leak ) of the leakage current measurement I leak which is delivered by the system 104, the result M(I keeper ) of the holding current measurement I keeper which is delivered by the system 122, and an evaluation E(I beam / I d ) of the thrust control current I beam / I d , respectively. The evaluation E(I beam / I d ) of the thrust control current can be directly the setpoint value which is prescribed by the manufacturer for this current in order to obtain the desired thrust force.Alternatively, the evaluation E(I beam / I d ) may be a result of a measurement of the thrust control current I beam / I d which is actually delivered by the interface 11 to the ion emission system 21. Such a measurement of the thrust control current I beam / I d may be carried out in real time in one of the ways available to those skilled in the art. The adder 121' thus provides as output an evaluation of the return current I body , denoted E(I body ), then the subtracter 120 calculates a difference Δ' which exists between the setpoint value I body_max_crit and this evaluation E(I body ). The operation of the control loop which is shown in [. Fig. 3 ] is then similar to that relating to [ Fig. 2], except that the holding current I keeper is now adjusted in real time during operation of the plasma thruster 2 so that the instantaneous values ​​of the return current I body remain substantially equal to the set value I boby_max_crit . For this, the difference Δ' which is calculated by the subtractor 120 is transmitted to the control input of the interface 12, and is used by the latter as an adjustment increment of the holding current I keeper to be delivered.

[0052] The embodiment of the invention which is illustrated by [ Fig. 4 ] is derived from that of [ Fig. 3] by replacing the evaluation E(I body ) of the return current I body which was provided by the adder 121' with a measurement of this return current. For this, a system 123 for measuring the return current I body is arranged on the electrical connection which connects the electron emitter 23 to the common return point CRP. This system 123 which is used to measure the return current I body can still be of a type known to those skilled in the art. The adder 121' is deleted. The principle of operation of the control loop which is shown in [ Fig. 4 ] is otherwise unchanged from that of [ Fig. 3], by replacing the evaluation E(I body ) of the return current I body which was provided by the adder 121' with a measurement result M(I body ) of this return current I body which is delivered by the system 123. Thus, the difference Δ" which is calculated by the subtracter 120 is transmitted to the control input of the interface 12, and used by the latter as an adjustment increment of the holding current I keeper to be delivered.

[0053] Generally, it may be advantageous for the holding current I keeper which is actually delivered to the holding electrode 22 to be more stable, with slower temporal variations. For this, this holding current I keeper may be determined by further applying a low-pass type transfer function, in particular of the proportional-integral type, to the deviations Δ, Δ' or Δ" which are transmitted as output by the subtractor 120 at successive instants of operation of the plasma thruster 2, for each of the embodiments of [ Fig. 2 ]-[ Fig. 4 ]. An optional low-pass filter 124 which is used for this purpose can then be inserted between the output of the subtractor 120 and the control input of the power supply interface 12 of the holding electrode 22.

[0054] It is understood that the invention may be reproduced by modifying secondary aspects of the embodiments that have been described in detail above, while retaining at least some of the advantages cited. In particular, some of the components used in these embodiments may be replaced by others with equivalent functions or which produce equivalent combinations of functions. Furthermore, some or all of the components used to adjust the holding current according to the invention may advantageously be incorporated into the power supply interface of the holding electrode, in order to obtain autonomous operation of this interface. Finally, the plasma propulsion system may incorporate any number of thrusters, with the associated electrical power supply interfaces.Optionally, plasma thrusters that are separate may share a common power supply interface while using the invention for that interface and those thrusters.

Claims

1. A power supply assembly (1) for a plasma thruster (2) of a spacecraft (100), adapted to transform an electrical power that is received from a power bus (101) of the spacecraft, into electrical voltages and currents that are supplied by said power supply assembly to the plasma thruster so that said plasma thruster generates an ion beam during an operation of the plasma thruster, the power supply assembly comprising a common return point (CRP) for collecting a return current (Ibody) from the plasma thruster during said operation, the plasma thruster (2) comprising an electron emitter (23) that is intended to neutralise the ion beam during the operation, and a keeper electrode (22) that is arranged at an output of the electrons emitted by the electron emitter, the electrical currents supplied by the power supply assembly (1) comprising at least one keeper current (Ikeeper) that is supplied to the keeper electrode and that constitutes a contribution to the return current (Ibody), the power supply assembly (1) being adapted to receive or produce successive evaluations of the return current (Ibody) or of a leakage current (Ileak) that flows from the common return point (CRP) to an electrical ground (102) of the spacecraft (100), the power supply assembly (1) being characterized in that it is arranged to adjust in real time during the operation, the keeper current (Ikeeper) depending on the evaluations of the return current (Ibody) or on the leakage current (Ileak), and in accordance with a current setpoint value (Ibody_max_crit; Iwarm_crit).

2. The power supply assembly (1) according to claim 1, wherein the current setpoint value (Ibody_max_crit; Iwarm_crit) that is used to adjust the keeper current (Ikeeper) in real time during the operation, relates to the return current (Ibody) or to a sum of the keeper current (Ikeeper) and of the leakage current (Ileak).

3. The power supply assembly (1) according to claim 1 or 2, arranged to adjust the keeper current (Ikeeper) in real time during the operation according to a feedback loop that comprises a subtractor (120), with a first input of the subtractor that is arranged to receive the current setpoint value (Ibody_max_crit; Iwarm_crit), and a second input of said subtractor that is arranged to receive the evaluations of the return current or to receive a sum of a measurement result of the keeper current and the evaluation of the leakage current, the power supply assembly being configured to adjust the keeper current depending on a difference result that is delivered as output by the subtractor (120).

4. The power supply assembly (1) according to claim 3, comprising a measuring system (104) that is arranged to output successive measurement results of the leakage current (Ileak), each measurement result of the leakage current constituting an evaluation of said leakage current, the power supply assembly (1) further comprising an adder (121) arranged to receive on two inputs of said adder, at each of a series of successive instants during the operation of the plasma thruster (2), the measurement result of the leakage current (Ileak) and the measurement result of the keeper current (Ikeeper) for said instant, and an output of the adder (121) being connected to the second input of the subtractor (120).

5. The power supply assembly (1) according to claim 3, comprising a measuring system (104) that is arranged to output successive measurement results of the leakage current(Ileak), each measurement result of the leakage current constituting an evaluation of said leakage current, the power supply assembly (1) further comprising an adder (121') arranged to receive on three inputs of said adder, at each of a series of successive instants during the operation of the plasma thruster (2), the measurement result of the leakage current (Ileak) and the measurement result of the keeper current (Ikeeper) for said instant, and also an evaluation of a thrust control current (Ibeam; Id) that is supplied at said instant by the power supply assembly to an ion emission system (21) of the plasma thruster (2), and an output of the adder (121') being connected to the second input of the subtractor (120).

6. The power supply assembly (1) according to claim 4 or 5, wherein the measuring system (104) that is arranged to output the successive measurement results of the leakage current (Ileak), is of a type adapted to carry out electric current measurements.

7. The power supply assembly (1) according to claim 4 or 5, further comprising a discharge conducting system (103) that is intended to electrically connect the common return point (CRP) to the electrical ground (102) of the spacecraft (100), and wherein the measuring system (104) that is arranged to output the successive measurement results of the leakage current (Ileak), is of a type adapted to carry out electrical voltage measurements, and is arranged to measure an electrical voltage that exists between terminals of at least one part of the discharge conducting system (103), and the power supply assembly (1) is configured to supply each measurement result of the leakage current (Ileak) as a result of a division of the electrical voltage measured between the terminals of the part of the discharge conducting system (103) by a resistance value of said part of the discharge conducting system.

8. The power supply assembly (1) according to claim 3, comprising a measuring system (123) arranged to output successive measurement results of the return current (Ibody), each measurement result of the return current constituting an evaluation of said return current, and wherein the power supply assembly (1) is arranged to transmit the successive measurement results of the return current (Ibody) to the second input of the subtractor (120) during the operation of the plasma thruster (2).

9. The power supply assembly (1) according to any one of claims 3 to 8, further configured to adjust the keeper current (Ikeeper) depending on the difference result that is outputted by the subtractor (120) in accordance with a proportional-integral transfer function applied to said difference result.

10. A plasma propulsion system (10) for a spacecraft (100), comprising: - at least one plasma thruster (2), said plasma thruster comprising an electron emitter (23) and a keeper electrode (22) that is arranged at an output of the electrons emitted by the electron emitter; and - a power supply assembly (1) that is in accordance with any one of the preceding claims, the plasma propulsion system (10) being arranged so that the power supply assembly (1) supplies to the keeper electrode (22) of the plasma thruster (2), during an operation of said plasma thruster, a keeper current (Ikeeper) that is adjusted in real time depending on the evaluations of the return current (Ibody) or on the leakage current (Ileak), in accordance with the current setpoint value (Ibody_max_crit; Iwarm_crit).

11. The plasma propulsion system (10) according to claim 10, wherein the plasma thruster (2) is of one of the following types: - a grid-type, continuous discharge or radio frequency ion thruster, wherein the electron emitter (23) comprises an electron gun that is intended to neutralise the ion beam during the operation of the plasma thruster, and the keeper electrode (22) is arranged at an electron output of the electron gun; or - a Hall effect thruster, wherein the electron emitter (23) is a neutralising hollow cathode of the Hall effect thruster, and the keeper electrode (22) is arranged at an output opening of the hollow cathode.

12. A spacecraft (100), comprising a plasma propulsion system (10) that is in accordance with claim 10 or 11.

13. A method for powering a plasma thruster (2) on board a spacecraft (100), comprising the following steps that are performed during a movement of the spacecraft in space: - supplying a keeper current (Ikeeper) to a keeper electrode (22) that is arranged at an electron output of an electron emitter (23) of the plasma thruster (2); - collecting successive evaluations of a return current (Ibody) that flows from the plasma thruster (2) to a common return point (CRP), or from a leakage current (Ileak) that flows from the common return point (CRP) to an electrical ground (102) of the spacecraft (100); - collecting a current setpoint value (Ibody_max_crit; Iwarm_crit); and - adjusting the keeper current (Ikeeper) that is supplied to the keeper electrode (22) depending on the evaluations of the return current (Ibody) or on the leakage current (Ileak), and in accordance with the current setpoint value (Ibody_max_crit; Iwarm_crit).

14. The method according to claim 13, wherein the keeper current (Ikeeper) that is supplied to the keeper electrode (22) is adjusted so that the return current (Ibody), or a sum of a measurement result of the keeper current and of the evaluation of the leakage current, corresponds to the current setpoint value (Ibody_max_crit; Iwarm_crit).

15. The method according to claim 13 or 14, wherein the current setpoint value (Ibody_max_crit; Iwarm_crit) is prescribed by a manufacturer of the plasma thruster (2).

16. The method according to any one of claims 13 to 15, implemented using a power supply assembly (1) that is in accordance with any one of claims 1 to 9, said power supply assembly being carried on board the spacecraft (100) and arranged to supply electrical voltages and currents to the plasma thruster (2), so that said plasma thruster generates an ion beam so as to apply a thrust force to the spacecraft.

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