FLIGHT TEST OF A GRATED IONIC PROPELLER

Abstract

A test method for a gridded ion thruster (2) is performed aboard a spacecraft (100) while said spacecraft is in outer space. The test method is a perveance test and allows for the determination of the thruster's operating point and the monitoring of its aging, including erosion of an acceleration grid (25) located at the outlet of a plasma chamber (20). (Shortcut figure: Figure 1a)

Description

Description Title of the invention: FLIGHT TEST OF A PROPELLER IONIC WITH GRIDS Technical field

[0001] — The present description relates to a method for flight testing an ion thruster with grids, as well as a power supply assembly for an ion thruster grids that are suitable for such a test. Prior art

[0002] — Grid ion thrusters are used for space vehicles such as satellites or space probes although their use is still to this day less common than that of Hall effect ion thrusters. These thrusters Grid ionic cells can be of the continuous discharge type, designated by the acronym GIT for “Gridded-Ion Thruster” in English and also called Kaufman type, or type RIT for "Radiofrequency Ion Thruster". These grid ion thrusters, RIT or GIT, each include at least: - a plasma enclosure; - a screen grille, which is located in front of or at the level of an outlet opening of the plasma enclosure; and - an acceleration grid, which is parallel to the screen grid and located on one side of this screen grid opposite the plasma speaker. During operation of such a thruster, positive ions are produced in the plasma enclosure pass through the screen grid and then the acceleration grid. For this, a power supply assembly, which is commonly referred to as PPU for "Power Processing Unit" in English, is also on board the space vehicle and arranged to, during operation of the thruster, simultaneously provide at least one positive electrical voltage to the screen grid and a negative electrical voltage to the acceleration grid, these positive and negative electrical voltages being determined by relative to a common reference node of the power supply assembly, which is commonly referred to as CRP for “Cathode Return Potential” in English, or sometimes NRP for “Neutralizer Return Potential” in English.

[0003] — The understanding and progressive improvement of ion thrusters with ne- grids require knowledge of the wear and aging phenomena that affect them pro- gradually when used on board space vehicles. For this, measurements and tests are necessary, to be carried out on board space vehicles, in order to characterize the state of wear and aging of each grid ion thruster throughout its use on mission. Indeed, the tests carried out on Earth are not not sufficiently representative of the environmental conditions to which each thruster is subjected in space, and it is necessary to collect measurement results throughout its use on missions. Such tests and measurements have the following aims in particular: - complete the general knowledge of grid ion thrusters, in particular concerning their aging in space environment conditions; - predict the future wear and / or aging of a grid ion thruster which is on mission on board a space vehicle, in particular to evaluate the lifespan of this thruster; and - characterize the state of wear and / or aging of a grid ion thruster which is on mission on board a space vehicle, in order to optimize the operating point at which this thruster is used, or to achieve a compromise between wear rate and intensity of the thrust force which is to be produced by the thruster. Erosion of the acceleration grid of a grid-type ion thruster is one of the main causes of its aging, which limits its lifetime. This erosion produces a progressive enlargement of the holes in the acceleration grid, and is caused by the bombardment of ions coming from the plasma enclosure, some of which hit the acceleration grid. For a fixed value of the negative electrical voltage applied to the acceleration grid, the state of erosion of this grid is revealed in particular by a modification of the back-circulation limit of the electrons generated by a neutralizer of the thruster. Document US 6,964,396 describes carrying out an EBS type test, for "Electron Back-Streaming" in English or measurement of electron back-flow, on board a space vehicle which is in extraterrestrial space. During such an EBS test, the negative electrical voltage which is applied to the acceleration grid is progressively decreased in absolute value, that is to say that this electrical voltage is varied from very negative values ​​to less negative values. Simultaneously, a beam current control loop, designated BCC for "Beam Current Control" in English, maintains a constant difference between an electrical current which is supplied to the screen grid and an electrical current which comes from the acceleration grid, by controlling a plasma generation electrical power which is delivered to the plasma enclosure of the thruster.This plasma generation electrical power is then measured for each value of the negative electrical voltage that is applied to the acceleration grid. Such an EBS test makes it possible to determine a value to be used for the negative electrical voltage that is applied to the acceleration grid, the absolute value of which is minimal to maximize the thrust produced at a given value of electrical power that is supplied to the screen grid, while reducing heating of the plasma enclosure and securing the operation of the thruster. Technical problem From this situation, an object of the present invention is to propose another test to be carried out while the space vehicle is in extraterrestrial space, to characterize the state of a grid ion thruster of this space vehicle, in particular the state of erosion of the acceleration grid of this thruster. Another object of the invention is to provide new information for managing the operation of a grid ion thruster which is on board a space vehicle, in particular such information which relates to the operation and lifetime of the thruster. Summary of the invention To achieve at least one of these aims or another, a first aspect of the invention provides a novel method for testing a gridded ion thruster on board a space vehicle, such as a satellite, a space probe, etc., this testing method being performed while the space vehicle is in extraterrestrial space. The thruster comprises at least the components that have been recalled previously: the plasma enclosure, the screen grid and the acceleration grid, the latter being on a side of the screen grid that is opposite the plasma enclosure so that, during operation of the thruster, positive ions that are produced in the plasma enclosure pass through the screen grid and then the acceleration grid.In particular, the thruster to which the test method of the invention is applied may be of one of the following types: continuous discharge grid ion thruster, as known under the designation GIT or also called Kaufman type, or radio-frequency discharge grid ion thruster, as known under the designation RIT. Possibly, the thruster to which the test method of the invention is applied may comprise at least one additional grid, in addition to the screen grid and the acceleration grid. Such a third grid is commonly called a deceleration grid. An electrical power supply assembly is also on board the space vehicle and arranged to, during operation of the thruster and during execution of the test method of the invention, simultaneously supply at least one positive electrical voltage to the screen grid and one negative electrical voltage to the acceleration grid, these electrical voltages being determined relative to the common reference node of the electrical power supply assembly. The method of the invention comprises the following steps: / 1 / determine at least respective values ​​of the positive electrical voltage and the negative electrical voltage for an identified operating point of the thruster which corresponds to its current operation; / 2 / provide two sweep limit values ​​for the positive electrical voltage, which are respectively greater than and less than the value of this positive electrical voltage for the identified operating point, or provide a start-of-scan limit value for the positive electrical voltage which is greater than the value of this positive electrical voltage for the identified operating point, and provide a maximum limit for an electrical current which flows from the accelerating grid to the electrical supply assembly, called the accelerating grid current;and / 3 / while the negative electrical voltage is kept constant and equal to its value for the identified operating point, varying the positive electrical voltage in a direction of variation which is constant between the scan limit values ​​provided in step / 2 / , by verifying that the acceleration grid current remains below the maximum limit provided, or decreasing the positive electrical voltage from the scan start limit value until the acceleration grid current reaches the maximum limit provided, and for each produced value of the positive electrical voltage, measuring the acceleration grid current and recording at least one measurement result of said acceleration grid current with the corresponding value of the positive electrical voltage. ; The maximum limit that is provided in step / 2 / for the accelerating grid current may depend on the beam current that leaves the thruster. In particular, it may be equal to the intensity of this beam current multiplied by a constant coefficient. Generally in the present description, the beam current that leaves the thruster is equal to a difference between an electric current that flows from the power supply assembly to the screen grid and the accelerating grid current, from which may be additionally subtracted a backflow current of electrons that are emitted by a neutralizer of the thruster and then collected by the screen grid. For most implementations of the invention, the maximum limit provided in step / 2 / for the accelerating grid current may be between 0.5% and 5% of the value of the beam current.For example, it can be approximately equal to 3.125% of the beam current value. The test method proposed by the invention is of the perveance test type. It makes it possible to identify the value of the positive electrical voltage which is applied to the screen grid, for which the acceleration grid current is minimal, that is to say for which a minimum quantity of ions which come from the plasma enclosure strike the acceleration grid. This value of the positive electrical voltage which minimizes the acceleration grid current corresponds to an operation of the thruster which reduces the erosion of its acceleration grid. The lifetime of the thruster can then be increased by adopting this operation, or its lifetime can be managed if other mission constraints require using an operating point of the thruster which does not correspond to the minimum of the acceleration grid current. The variation range of the positive electrical voltage which is implemented in step / 3 / , being limited on the side of the lower values ​​of this positive electrical voltage either by one of the scanning limit values ​​provided in step / 2 / , or because the maximum limit for the acceleration grid current is reached, depends on the amount of information which is desired on the thruster. An advantage of the method of the invention is that it allows measurements to be taken regularly during operation of the thruster that provides thrust to the space vehicle, to detect the erosion state of its grids, and to determine whether its operating point needs to be corrected. The thrust force that is produced by the thruster is not interrupted by the execution of the test method of the invention. Another advantage of the method of the invention is that it allows the operation of the thruster to be adjusted according to its aging state, without damaging the thruster, in a way that is particularly safe and robust so that it can be carried out automatically and in a space environment. Yet another advantage of the method of the invention is that it allows the space vehicle to be tested in flight, and multiple operating parameters of the thruster to be transmitted remotely in order to understand its behavior under operational conditions. The present invention therefore makes it possible to improve the understanding of the evolution of the characteristics, in particular electrical, thermal or material-related, of a grid ion thruster which is used on board a space vehicle throughout the lifetime of this space vehicle. It in fact provides measurements of certain characteristics of the thruster during the mission of the space vehicle. Finally, yet another advantage of the invention lies in the superior precision of the tests and measurements that are carried out while the space vehicle is in extraterrestrial space, compared to tests and measurements carried out on Earth. Indeed, many characteristics of the thruster have very different values ​​on Earth compared to their actual values ​​in extraterrestrial space, in particular because of the lack of representativeness of the environmental parameters, e.g. pressure, narrowness of the vacuum chambers, etc., which are used on Earth to simulate the vacuum of space. Thus, the test method of the invention provides more accurate knowledge of the behavior of grid ion thrusters in extraterrestrial space throughout their operation. This knowledge makes it possible in particular to better adjust the electrical parameters of a grid ion thruster, i.e. electrical currents and voltages, and also fluidic parameters, i.e.flow rates and pressures of the gases used in this thruster. Such an improved adjustment in turn allows for optimized exploitation of the thruster grids, in particular by minimizing their erosion. Advantageously, the positive electrical voltage can be varied in step / 3 / automatically tically by the power supply assembly, in accordance with a programming of this power supply assembly. Possibly, and in order that results of execution of the test method of the invention can be studied and / or exploited on Earth, the method may further comprise transmitting from the space vehicle to a station which is located on Earth, at least some of the measurement results and values ​​recorded in step / 3 / . Such transmission may be included in a telemetry procedure which is used for the space vehicle. In certain implementations of the invention, particularly those for the purpose of reducing acceleration grid erosion, the method may further comprise the following step: / 4 / determine a minimum value of the acceleration grid current from among the measurement results recorded in step / 3 / for this acceleration grid current, as well as the respective corresponding values ​​of the positive electrical voltage and the negative electrical voltage. When step / 4 / is executed, the method may also further comprise: - updating the identified operating point of the thruster in accordance with the respective values ​​of the positive electrical voltage and the negative electrical voltage which correspond to the minimum value of the accelerating grid current, as determined in step / 4 / , possibly with a predetermined margin of deviation between the value of the positive electrical voltage which was determined for the minimum value of the accelerating grid current and the value of the positive electrical voltage of the updated identified operating point; then - activate a new thruster operation that is consistent with the updated identified operating point. Such updating and activation of the thruster operating point can be controlled remotely from the station which is located on Earth, or performed automatically on board the space vehicle. Also when step / 4 / is executed, and when the positive electrical voltage has been decreased in step / 3 / from the scan start limit value until the accelerating grid current reaches the provided maximum limit, the method may further comprise the following steps: / 5 / determining a threshold value for the positive electrical voltage which corresponds to a bend in a curve of the measured values ​​of the acceleration grid current as a function of the values ​​of the positive electrical voltage, on the side of the lowest values ​​of this positive electrical voltage; and / 6 / calculate a difference between the value of the positive electrical voltage which corresponds to the minimum value of the acceleration grid current, as it has been determined in step / 4 / , and the threshold value determined in step / 5 / for the positive electrical voltage. The difference result calculated in step / 6 / is a measure of the erosion state of the accelerator grid, and therefore constitutes a measure of the aging of the thruster. The sequence of steps / 1 / to / 6 / may be repeated several times, each time at a different time while the spacecraft is in extraterrestrial space, and then a new value is determined for the difference between the value of the positive electrical voltage which corresponds to the minimum value of the accelerator grid current, and the threshold value of this positive electrical voltage, by extrapolation of the difference values ​​calculated at each execution of the sequence of steps / 1 / to / 6 / .Indeed, these difference values ​​calculated at each execution of the sequence of steps / 1 / to / 6 / characterize an erosion profile of the acceleration grid, as a function of the operating time of the thruster, and this time profile can be extended by extrapolation to provide a forecast of the difference value which is relative to a future date. The power supply assembly may comprise a device for controlling a plasma generation electrical power that is delivered to the plasma enclosure by this power supply assembly, this control device being designed to maintain a constant value of difference between the electrical current that flows from the power supply assembly to the screen grid and the acceleration grid current. Such a control device corresponds to the beam current control loop, or BCC for "Beam Current Control" in English. Then, in first possible implementations of the test method of the invention, step / 3 / may be executed while the value of the difference between the electrical current that flows from the power supply assembly to the screen grid and the acceleration grid current is maintained constant by the control device.This difference value is equal to the beam current in the absence of additional current called backflow. Alternatively, in other possible implementations of the test method of the invention, step / 3 / may be executed while the plasma generation electrical power that is delivered to the plasma enclosure by the electrical power supply assembly is kept constant. For this purpose, and if present, the plasma generation electrical power control device is kept deactivated while step / 3 / is executed. The value of the difference between the electrical current that flows from the electrical power supply assembly to the screen grid and the acceleration grid current is then no longer constant, because of the profile adopted by the acceleration grid current under the effect of the variation of the positive electrical voltage that is supplied to the screen grid. A second aspect of the invention provides a power supply assembly for a gridded ion thruster, the thruster comprising: a plasma enclosure, a screen grid which is located in front of or at an outlet opening of the plasma enclosure, and an accelerating grid which is parallel to the screen grid and located on a side thereof which is opposite the plasma enclosure, the power supply assembly comprising: - the common reference node: - a positive electrical power supply unit, which is intended to be connected to the screen grid of the thruster to provide this screen grid, during operation of the thruster, with a positive electrical voltage relative to the common reference node; - a negative electrical power supply unit, which is intended to be connected to the acceleration grid of the thruster to provide this acceleration grid, during operation of the thruster, with a negative electrical voltage relative to the common reference node: - a variation module, which is arranged to vary, preferably automatically, the positive electrical voltage supplied by the positive electrical power supply unit; - measuring and recording modules, arranged to measure and record values ​​of the acceleration grid current; and - a controller, which is configured to activate the variation module and the measurement and recording modules, so that variable values ​​are successively produced for the positive electrical voltage while the negative electrical voltage is kept constant, and that a value of the acceleration grid current is measured and recorded for at least one supplied value of the positive electrical voltage, when the electrical power supply assembly is connected to the thruster to enable operation of this thruster. Such a power supply assembly is suitable for performing a test method which is in accordance with the first aspect of the invention. Each of the variation, measurement and recording modules of the power supply assembly may be of the software or recorded program type, designated by "software" in English, or of the hardware type, designated by "hardware". Possibly, the power supply assembly may further comprise: - a device for controlling the plasma generation electrical power delivered to the plasma enclosure, designed to maintain the beam current constant, and as designated by beam current control loop, or BCC. A third aspect of the invention provides a plasma propulsion system which comprises: - the grid ion thruster; and - a power supply assembly which is in accordance with the second aspect of the invention, and which is connected to the thruster to produce operation of the latter. Finally, a fourth aspect of the invention provides a space vehicle which comprises a plasma propulsion system in accordance with the third aspect. Brief description of the figures 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.1a] shows, in a schematic and simplified manner, a space vehicle which is equipped with a GIT type grid ion thruster, and on board which the test method of the invention can be used; [Fig.1b] corresponds to [Fig.1a] for a RIT-type grid ion thruster; [Fig.2a] is an example of a diagram presenting results which are obtained by the test method of the invention, according to a first possible sequence of the method; [Fig.2b] corresponds to [Fig.2a] for a second possible sequence of the process; [Fig.3a] illustrates a situation of over-perception; [Fig.3b] illustrates an optimal situation of perveance; and [Fig.3c] illustrates a situation of under-perception. Detailed description of the invention In these figures, all elements are represented only symbolically, and identical references indicated in different figures designate identical elements or which have identical functions. In [Fig.1a] and [Fig.1b], the reference 100 designates a space vehicle, regardless of the type of this vehicle, for example a satellite or a space probe. Inside this space vehicle 100, the reference 10 designates 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 space vehicle 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 grid ion thruster 2.The power supply assembly 1 and the grid ion thruster 2 which are shown in each of these two figures are associated with each other so that the thruster 2 is supplied with electrical energy in an appropriate manner by the power supply assembly 1, in order to produce the thrust force which is desired at each instant of operation of the plasma thruster, while the space vehicle 100 is in . extraterrestrial space. Typically, the power supply assembly 1 is connected between an electrical power bus and an electrical ground (not shown) of the space vehicle 100, while the various components of the thruster 2 are supplied with electrical currents and voltages via the electrical power supply assembly 1. During operation of each of the grid ion thrusters of [Fig.1a] and [Fig.1b], the thrust force which is produced by this thruster and applied to the space vehicle 100 results from the production, by the grid ion thruster, of a beam of ions which is denoted ions in the figures. For the sake of clarity of the present description, the components of the plasma propulsion system 10 that are shown in [Fig. 1a] and [Fig. 1b] are limited to those covered by the invention. In particular, each of the grid ion thrusters 2 of [Fig. 1a] and [Fig. 1b] comprises one or more neutralizer power supply units, as well as a gas management system, which are not shown, and the power supply assembly 1 comprises additional power supply units and corresponding interfaces, which are also not shown. For the space vehicle 100 of [Fig.1a], the gridded ion thruster 2 is of the continuous discharge type designated by GIT for "Gridded Ion Thruster", or called Kaufman type. It comprises a plasma enclosure 20, which is provided with an anode 21 and a cathode 22. The anode 21 may be located around an outlet opening of the plasma enclosure 20, through which the ions are intended to exit during operation of the thruster 2. The cathode 22 may be located at the bottom of the plasma enclosure 20, opposite its outlet opening. In a known manner, the plasma enclosure 20 is provided with an electromagnet 23, commonly called a "magnet", which helps to confine the plasma inside the enclosure 20.To produce the plasma in the enclosure 20, and thus constitute a source of the ions which are ejected towards the outside of the vehicle 100, a continuous electric discharge is created in the enclosure 20 between the anode 21 and the cathode 22, at the same time as an ionizable gas is introduced therein. The grid ion thruster 2 further comprises at least two electrically conductive grids, which are arranged parallel and at a distance from each other in front of the outlet opening of the plasma enclosure 20. The first grid, called the screen grid and designated by the reference 24, has the main functions of controlling the quantity of ions which leave the plasma enclosure 20 and of participating in accelerating the ions which leave the plasma enclosure 20. For this, the screen grid 24 is brought to a positive electrical voltage which is denoted Vpyy. The second grid, called the acceleration grid and designated by the reference 25, contributes to accelerating the ions which leave the plasma enclosure 20, in cooperation with the screen grid 24, and creates an electrical potential barrier between the screen grid and the neutralizer 27 described later, for electrons emitted by the latter. For this, the acceleration grid 25 is brought to a negative electrical voltage which is noted Vyyy. The screen grid 24 is intermediate between the outlet opening of the plasma enclosure 20 and the acceleration grid 25. Possibly, the thruster 2 can comprise at least a third additional grid. Such a third grid, called a deceleration grid, can be electrically connected to the electrical ground of the space vehicle 100. The power supply assembly 1 comprises a first power supply unit which is dedicated to the power supply of the plasma enclosure 20, designated by the reference 11 and noted DC-plasma, a second power supply unit which is dedicated to the polarization of the screen grid 24, designated by the reference 12 and noted PHV for “positive high voltage” in English, and a third power supply unit which is dedicated to the polarization of the acceleration grid 25, designated by the reference 13 and noted NHV for “negative high voltage”. In the general part of the present description, the power supply unit 12 has been called the positive power supply unit, and the power supply unit 13 has been called the negative power supply unit. The power supply assembly 1 has a common reference node which is commonly referred to as CRP for “Cathode Return Potential” in English, or sometimes NRP for “Neutralizer Return Potential” in English.This common reference node CRP is itself electrically connected to the electrical ground of the spacecraft 100 by a charge conductive system 19, sometimes called a "bleed resistor" in English. The positive Vpyy and negative Vyuy electrical voltages are defined relative to this common reference node CRP. For this, the power supply unit 12, which is dedicated to the polarization of the screen grid 24, has a positive output terminal which is electrically connected to this screen grid 24, and a negative output terminal which is electrically connected to the common reference node CRP. Thus, the electrical voltage Vpyy which is applied to the screen grid 24 is positive. The electrical current which flows through the positive output terminal of the power supply unit 12 towards the screen grid 24 is noted Ipuv, and called the screen grid current.Similarly but reversing the polarity, the power supply unit 13, which is dedicated to the polarization of the acceleration grid 25, has a negative output terminal which is electrically connected to this acceleration grid 25, and a positive output terminal which is electrically connected to the common reference node CRP. The electric current which flows through the negative output terminal of the power supply unit 13 from the acceleration grid 25 is denoted Ivuv, and called the acceleration grid current. The electric voltage Vxuv which is thus applied to the acceleration grid 25 is negative. The power supply unit 11, which is dedicated to powering the plasma enclosure 20, has a positive output terminal which is electrically connected to the anode 21, and a negative output terminal which is connected to the common reference node CRP. The electric current which flows through the positive output terminal of the power supply unit 11 towards the anode 21 is denoted I, and called the plasma discharge current. The cathode 22 is electrically connected to the common reference node CRP. The electric power for generating the plasma, which is supplied by the power supply unit 11 to the plasma enclosure 20, through the anode 21 and the cathode 22, is equal to the product of the value of the discharge current I, by the value of an electric voltage V4 of the anode 21, determined with respect to the common reference node CRP. This plasma generation power can be determined internally to the power supply unit 11. The grid ion thruster 2 further comprises a neutralizer 27, denoted NEUTR, whose function is to emit electrons to the outside of the thruster 2 in order to neutralize the ions which are emitted to produce the thrust. In a known manner, this neutralizer 27 is electrically powered by several dedicated power supply units which are integrated into the electrical power supply assembly 1, but without being shown in the figures. A dedicated electrical connection connects the neutralizer 27 to the common reference node CRP to conduct to this node a neutralizer return current which is denoted Iseurr. The neutralizer 27 thus determines the electrical potential of the common reference node CRP.In general, the beam current, denoted Iseams, is given by the formula: Tbeum = Iruv - Ixuv - Ipuv_xss, WHERE Ipuv_eps @St An additional current which is supplied by the power supply unit 12 to the screen grid 14, without participating in the thrust which is produced because it is consumed by the back-circulation of electrons from the neutralizer 27 to the screen grid 14. The beam current Team €St the electric current which corresponds to the beam of ions which are emitted by the thruster 2 towards the outside of the space vehicle 100. During nominal operation of the thruster 2, that is to say during operation for which the back-circulation of electrons is very low and negligible, the beam current Team is substantially equal to the difference between the screen grid current and the acceleration grid current: Team % Ipuv - Intv- . The other references which are indicated in [Fig.1a], and which will possibly be repeated in [Fig.1b], will be described later. For the space vehicle 100 of [Fig. 1b], the grid ion thruster 2 is of the radiofrequency discharge type designated by RIT for "Radiofrequency Ion Thruster". It still comprises a plasma enclosure 20, but this is provided with a coil 26 which surrounds a side wall of this enclosure to generate a radiofrequency electromagnetic field inside the plasma enclosure 20. The radiofrequency electromagnetic field produces the ionization of the gas which is introduced into the plasma enclosure 20, and which constitutes the source of ions. For such a mode of plasma generation, the coil 26 is supplied with radiofrequency current, denoted Igr, by a generator 3, denoted RFG for "radiofrequency generator", which can also be integrated in the electrical power supply assembly 1. This RFG generator 3 is itself supplied by a dedicated power supply unit of the electrical power supply assembly 1, designated by the reference 14 and called RF ionization stage driver, or "RF ionization stage driver" in English. This power supply unit 14 is denoted RF-driver in [Fig.1b] and replaces the power supply unit 11 of [Fig.1a] in the electrical power supply assembly 1. The electrical power for plasma generation in such a case of RIT thruster is calculated from the value of the radiofrequency current Ir and the value of the output electrical voltage Vpr of the RFG generator 3. The arrangement of the screen grid 24 and the acceleration grid 25, and their respective functions with their dedicated power units 12 and 13, as described above with reference to [Fig.1a], apply identically to the RIT type gridded ion thruster of [Fig.1b]. The electrical power supply assembly 1, for both types of GIT and RIT thrusters, further comprises a variation module 15, denoted VAR., for controlling the power supply unit 12, in order to vary the positive electrical voltage Vpuy which is produced by the latter. It also comprises measurement and recording modules, which are designated overall by the reference 16, in order to measure and record the values ​​of the acceleration grid current Ixyv. Finally, a controller 17, denoted CTRL and internal to the electrical power supply assembly 1, controls executions of a perveance test method, as now described. The permeability test method of the present invention is executed when the space vehicle 100 is in flight in extraterrestrial space, from an ongoing operation of the thruster 2, GIT or RIT, electrically powered by the assembly 1. This operation is controlled by the controller 17 by imposing several output values ​​of the electrical power supply assembly 1, including those of the positive Vpuy and negative Vyuv electrical voltages. These output values ​​of the electrical power supply assembly 1 which are thus produced participate in determining the operating point of the thruster 2. Each operating point which is available for the thruster 2 is identified, in order to distinguish it from other operating points in the case where several are possible. The operating point also includes the setting by the controller 17 of an output value of the power supply unit 11 or 14, depending on the GIT or RIT type of the thruster 2. For the GIT case of [Fig.1a], this output value is that of the plasma generation power which is supplied by the power supply unit 11 to the plasma enclosure 20. Two operating modes are possible alternatively: with beam current Ipeam WHICH is constant, or with generation power of plasma which is constant. For this, a servo device is associated with the power supply unit 11, which is adapted to vary in real time the plasma generation power so as to maintain the beam current Ivrum CONStant during the current operation of the thruster 2. This servo device, called BCC for "beam current control", can comprise a module 18a which is arranged to measure the current Ipyy which is transmitted by the power supply unit 12 to the screen grid 24, and a module 18b which is adapted to control the electrical power which is supplied at the output by the power supply unit 11. The module 18a transmits to the module 18b the value of the current Ipyy, at the same time as the value of the current Ivuy which returns from the acceleration grid 25 to the power supply unit 13 is also transmitted to the module 18b by the measurement and recording module 16.The operating mode of the thruster 2 for which the beam current I, €St constant, is obtained by activating the BCC servo device, and the operating mode for which the plasma generation power is constant, is obtained by deactivating the BCC servo device. For the RIT case of [Fig. 1b], the BCC servo device still comprises the module 18b, but arranged to control the electrical power which is supplied by the power supply unit 14 to the RFG generator 3. The current measurement modules 16 and 18a are arranged for the RIT case as for the GIT case, respectively at the outputs of the power supply units 12 and 13. When the BCC servo device is activated, for the RIT case, the plasma generation power which is supplied by the RFG generator 3 to the coil 26 is adjusted in real time to maintain the beam current Is:m CONStant.Again in the RIT case, when the BCC servo device is deactivated, the coil 26 is supplied from the power supply unit 14, via the RFG generator 3, with a plasma generation power which is kept constant. From the operating point which is in use for the thruster 2, whether it is GIT or RIT, and to execute the perveance test, the controller 17 commands to keep constant the negative electrical voltage Vnuyy which is applied to the acceleration grid 25, and to vary, via the module 15, the positive electrical voltage Vpuy which is applied to the screen grid 24. For each value of the positive electrical voltage Vpuy which is thus produced, the module 16 measures and records the value of the acceleration grid current Iysv. Such a perveance test can be carried out alternatively while the servo device BCC is activated, or deactivated, in both GIT and RIT cases.When performed while the BCC servo device is deactivated, it may be necessary to turn off the thruster 2 after the end of the perveance test, then restart it after the BCC servo device is reactivated to begin a new thrust production duration. The perveance test is preferably performed automatically. by the power supply assembly 1, in accordance with a programming of its controller 17. [Fig.2a] and [Fig.2b] illustrate two possible sequences of such a pervasiveness test. The horizontal axis of the diagrams in these figures marks the time, denoted t, the left vertical axis marks the values ​​of the positive electrical voltage Vpyy which are successively controlled by the controller 17, and the right vertical axis marks the values ​​of the acceleration grid current Inyy which are measured by the module 16 for the controlled values ​​of the voltage Vpyy, as results of the test. The test can start from a scan limit value Vpuv_,p which is greater than the value of the positive electrical voltage Vpyy which corresponds to the operating point in use for the thruster 2, then the positive electrical voltage Vpyy is gradually reduced during the test from this value Vpyv_,up.This reduction of the voltage Vpyy is carried out by successive descending steps, with a decrement AVpuy and an individual duration of step At which are selected to establish a compromise between the duration of the test and its precision. The end of the sweep of the values ​​of the positive electric voltage Vpuv can be determined in two alternative ways: either by another value of sweep limit Vpuv_ins which is lower than the value of the positive electric voltage Vpyy which corresponds to the operating point in use for the thruster 2, or by a maximum limit relating to the acceleration grid current Ivy which is measured.Indeed, in order not to degrade the thruster 2, the acceleration grid current Inuv must remain lower than a fraction of the beam current Isa, for example lower than 0.03125-l,zum- During such a downward variation of the positive electrical voltage Vpuy, the acceleration grid current Iyuy begins by decreasing progressively, reaching a minimum value Inuv_min Then increases, first slowly and then more quickly while the positive electrical voltage Vpuy continues to be reduced with a reduction speed which is kept constant.The variation curve of the acceleration grid current Ivay thus has the following two characteristics: the minimum value Ivuv_min of the acceleration grid current Ivy which is obtained for the value Vpuv_min of the positive electrical voltage Vpuy, and the rapid change of slope in the curve of the values ​​of the acceleration grid current Inyv, occurring at the value V puv_change of the positive electrical voltage Vpyv. This last value Vprv_change which is associated with the rapid change of slope in the curve of the values ​​of the acceleration grid current Ivuv, has been called the threshold value of the positive electrical voltage in the general part of this description. The value Vpuv_change, or preferably the value of the difference Vpuv_min - VPHY_change» constitutes a characterization of the state of erosion of the grids 24 and 25, mainly of the acceleration grid 25. [Fig.2a] corresponds to a progression of the persistence test for which the progressive reduction. of the positive electrical voltage Vpyy is stopped when the acceleration grid current Ivuy reaches the maximum limit set for this current, i.e. 0.03125-L,cam in the example considered. Such a sequence provides the value Vpuv_change . [Fig.2b] corresponds to an alternative sequence of the perveance test for which the progressive reduction of the positive electrical voltage Vpuy is stopped when this voltage Vpyy reaches the scanning limit Vpuv ing In this last case, the scanning limit Vpyv._inr is greater than the value of the positive electrical voltage Vpyy for which the acceleration grid current Inuy would have reached its maximum limit set. The results of the perveance test, that is to say, the value of the negative electrical voltage Vyuy for which it was carried out, and the successive associated values ​​of the positive electrical voltage Vpyy and of the acceleration grid current Ivyv, can be used on board the space vehicle 100 or transmitted to a ground station under the control of an on-board computer of the space vehicle 100, designated by the reference 110 and noted OBC for “on-board computer” in English in [Fig. 1a] and [Fig. 1b]. [Fig.3a]-[Fig.3c] schematically illustrate different perveance situations for the exit of ions from the plasma enclosure 20. Each of these figures is a section of the exit opening of the plasma enclosure 20, with a hole of the screen grid 24 and, opposite it, a hole of the acceleration grid 25. For the illustrated case, the hole of the acceleration grid 25 has an opening surface which is smaller than that of the hole of the screen grid 24. The positive electrical voltage Vpyv which is applied to the screen grid 24 determines the shape of a sheath of the plasma which is contained in the enclosure 20, at the level of the hole of this screen grid 24. This sheath of the plasma is designated by the reference PI_Sh, for “plasma shealth” in English. The situation shown by [Fig.3a] corresponds to an over-permeance, for which the plasma sheath PI_Sh is substantially flat in the opening of the hole of the screen grid 24.The ions that pass through this opening in the screen grid 24 have substantially parallel trajectories, so that a peripheral part of these ions strike the acceleration grid 25 on the edges of the hole of the latter, producing a significant value for the acceleration grid current Ix;v. Because of this ion bombardment, such an over-perveance situation causes a progressive erosion of the acceleration grid 25. Conversely, the situation that is shown by [Fig.3c] corresponds to an under-perveance, for which the plasma sheath PI_Sh is too concave in the opening of the hole of the screen grid 24. The ions that pass through this opening in the screen grid 24 then have trajectories inclined towards the median axis of the opening, and all the more inclined as the ions pass through the opening of the hole close to its peripheral edge.Then a part of the ions which come from the periphery of the hole in the screen grid 24 still hits the edges of the hole in the acceleration grid 25, now after having crossed the median axis. A significant value. This again results in the acceleration grid current Iyyy. Such an under-perveance situation therefore also causes erosion of the acceleration grid 25. Finally, [Fig.3b] shows an intermediate situation, for which the plasma sheath PI_Sh is slightly concave in the opening of the hole in the screen grid 24. The ions that pass through this opening then have trajectories that gently narrow and pass through the opening of the hole in the acceleration grid 25 without hitting the latter, or intercepting it minimally. The perveance situation of [Fig.3b] is optimal in producing erosion of the acceleration grid 25 that is minimal, and corresponds to the minimum value Ixuy min of the acceleration grid current Ivuv. Again with reference to the perveance test sequences illustrated by [Fig.2a] and [Fig.2b], the first part of the reduction of the positive electrical voltage, corresponding to Vpyy larger than Vpv_min: corresponds to under-perveance situations as shown by [Fig.3c], and the last part of the reduction of the positive electrical voltage, corresponding to Vzuy smaller than Vpuv_min, corresponds to over-perveance situations as shown by [Fig.3a]. The value Vpuv_nin produces the optimal perveance situation, as shown by [Fig.3b]. A first possible use of the results of the perveance test is to update the operating point of the thruster 2. The updated operating point has the value of the negative electrical voltage Vnuv for which the test was carried out, and a value close to Vpzv_nin €St assigned to the positive electrical voltage Vp nv. The updated value for the positive electrical voltage Vpyy can take into account a predetermined safety margin, so that this value of Vpyy for the updated operating point is sufficiently far from the threshold value Vpiv_change- This results in a possible margin of deviation between the value of Vpuy used for the update of the operating point and the value Vpuv_min. In this way, the erosion of the acceleration grid 25 is reduced for the future operation of the thruster 2 according to this operating point.Such an update can be performed either automatically on board the spacecraft 100, by the controller 17, or remotely controlled by an operator after the test results have been transmitted to him. The operation of the thruster 2 can then be continued after the perpetuity test in accordance with the updated operating point. A second possible use of the results of the perveance test consists in monitoring the evolution of the aging of the propellant 2. Indeed, the value Vpuv_change corresponds to the limit of effectiveness of the screen grid 2 to control the quantity of ions which leave the plasma enclosure 20. Because of the aging of the propellant 2, the values ​​VpHv_change which are obtained by perveance tests carried out on different dates increase with the duration of use of the propellant 2, reducing the value of the difference Vpuv_min - VPHY_change- The curve of these difference values ​​Vpuv_min - V PHY_change as a function of the thruster's usage time makes it possible to predict its aging, and possibly to preferentially adopt operating points which cause slower aging of the thruster. Furthermore, it is also possible to implement a pervasiveness test which is still in accordance with the invention, but which is limited to a restricted interval of variation of the positive electrical voltage Vpyy around its value for the operating point in use. Such a test is shorter, but still makes it possible to determine the value Vpuv_nin €t to update the operating point accordingly. For such implementations, the positive electrical voltage Vpyy is varied between two predetermined limit values ​​Vpuv sup €t Vpuv_ns of the scan, in the increasing direction or in the decreasing direction, while verifying that the acceleration grid current Ivuv remains constantly below the maximum limit set, for example below 0.03125-Leum- It is understood that the invention may be reproduced by modifying secondary aspects of the embodiments which have been described in detail above, while retaining at least some of the advantages cited. In particular, all the numerical values ​​which have been cited have been cited only for illustration purposes, and may be changed depending on the application considered.

Claims

Claims

1. Method for testing a grid ion thruster (2) on board a space vehicle (100), the method being performed while the vehicle spacecraft is in extraterrestrial space, the propellant comprising at LESS! - a plasma enclosure (20); - a screen grid (24), which is located in front of or at a level plasma enclosure outlet opening (20); and - an acceleration grid (25), which is parallel to the screen grid (24) and located on one side of said screen grid opposite the plasma enclosure (20), so that, during operation of the propellant (2), ions positives which are produced in the plasma enclosure (20) pass through the screen grid (24) then the acceleration grid (25), an electrical power supply assembly (1) also being on board the vehicle spatial (100) and arranged for, during operation of the thruster (2) and during an execution of the test method, simultaneously provide at least a positive electrical voltage (Vpuv) to the screen grid (24) and a negative electrical voltage (Vxuv) to the acceleration grid (25), the positive and negative electrical voltages being determined with respect to a common reference node (CRP) of the power supply assembly electric, and the method being characterized in that it comprises the following steps: / 1 / determine at least the respective values ​​of the electrical voltage positive (Vpuy) and negative electrical voltage (Vxuv) for a point identified operating mode of the propellant (2) which corresponds to a func- current operation of said propellant; / 2 / provide two sweep limit values ​​for the electrical voltage positive (Vpuv), which are respectively greater than and less than the value of said positive electrical voltage for the operating point- identified operation, or provide a start limit value of sweep (Vpuy_sup) FOR the positive electrical voltage which is higher to the value of said positive electrical voltage for the operating point- identified operation, and provide a maximum limit for a current electric current flowing from the acceleration grid (25) to the assembly power supply (1), called accelerating grid current (I suv): and / 3 / while the negative electrical voltage (Vxuy) is kept constant and equal to the value of said negative electrical voltage for the point of identified operation, vary the positive electrical voltage (Vpuv) according to a direction of variation which is constant between the limit values ​​of sweep provided in step / 2 / , checking that the grid current acceleration (Iyuy) remains below the maximum limit provided, or decrease the positive electrical voltage from the limit value of start of sweep (Vpuv_sup) UNTIL the grid current acceleration reaches the maximum limit provided, and for each value produced positive electrical voltage, measure the grid current acceleration and record at least one measurement result of said acceleration grid current with the corresponding value of the positive electrical voltage.

2. A method according to claim |, wherein the electrical voltage positive (Vpuy) is varied in step / 3 / automatically by the set power supply (1), in accordance with a programming of said power supply assembly.

3. A method according to claim 1 or 2, further comprising transmit from the space vehicle (100) to a station located on Earth, at least some of the measurement results and values ​​recorded at step / 3 / . |Claim 4] A method according to any preceding claim, further comprising the following step: / 4 / determine a minimum value (Ivuv_min) of the grid current acceleration (Ivy) among the measurement results recorded at step / 3 / for the said acceleration grid current, as well as the cor- respective corresponding (Vpxv_mins Vnuy) of the electric voltage positive (Vpyy) and negative electrical voltage (Vnuv)-

5. The method of claim 4, further comprising: - update the identified operating point of the thruster (2) according to the respective values ​​(Vpuv_mins Vnuv) of the voltage positive electrical voltage (Vpyy) and negative electrical voltage (Vwuv) which correspond to the minimum value (Ivnv_min) of the grid current acceleration (Ivyv), as determined in step / 4 / , possibly with a predetermined margin of deviation between the voltage value positive electric current determined for said minimum value of the current of acceleration grid and the value of the positive electrical voltage of the updated identified operating point; then - activate a new operation of the thruster (2) which is compliant at the updated identified operating point.

6. A method according to claim 4 or 5, wherein the voltage positive electrical current (Vpuy) is decreased in step / 3 / from the value of scan start limit (Vpuy_sup) UNTIL the current of acceleration grid (Içuy) reaches the maximum limit provided, and the method further comprises the following steps: / 5 / determine a threshold value (Vpuy_change) FOR the electrical voltage positive (Vpuy) which corresponds to a bend in a curve of values measured acceleration grid current (Ivuy) as a function of the values ​​of positive electrical voltage, on the side of the highest values low levels of said positive electrical voltage; and / 6 / calculate a difference between the value (Vpuv_min) of the voltage positive electrical (Vpuy) which corresponds to the minimum value (Ivuv_min) of the accelerating grid current (Inuv), as determined in step / 4 / , the threshold value (Vprv_change) determined in step / 5 / for said positive electrical voltage.

7. A method according to claim 6, wherein the sequence of steps / 1 / to / 6 / is repeated several times, each time at a different time while the space vehicle (100) is in extraterrestrial space, then a new value is determined for the difference between the value (Vpuv_min) of the positive electrical voltage (Vpuy) which corresponds to the minimum value (Inuv_min) of the acceleration grid current (Inuv) and the threshold value {(Vpuv_change) of said positive electrical voltage, by extrapolation of the difference values ​​calculated at each execution of the sequence of steps / 1 / to / 6 / .

8. A method according to any one of claims 1 to 7, wherein the power supply assembly (1) comprises a device control (18a, 18b) of an electric power generation of plasma which is delivered to the plasma enclosure (20) by said assembly power supply, said servo device being designed to maintain constant a difference value between a current electric (Ipuy) which circulates from the power supply assembly to the screen grid (24) and the acceleration grid current (Ivuv), and according to which step / 3 / is executed while the value of the difference between the electric current (Iruy) which flows from the ali- electrical supply (1) to the screen grid (24) and the grid current acceleration (Ivpv) is kept constant by the device of enslavement (18a, 18b).

9. A method according to any one of claims 1 to 7, wherein step / 3 / is executed while an electrical power of ge- plasma generation which is delivered to the plasma enclosure (20) by the power supply assembly is kept constant.

10. A method according to any preceding claim, which the maximum limit provided in step / 2 / for the grid current acceleration (Ivuy) is between 0.5% and 5% of a value of a beam current (Ipeum) which comes out of the thruster (2).

11. A method according to claim 10, wherein the maximum limit provided in step / 2 / for the acceleration grid current (Ixuvy) is substantially equal to 3.125% of the beam current value (Ieam)-

12. Method according to any one of the preceding claims, following which the propellant (2) is of one of the following types: ion propellant continuous discharge grid, or continuous discharge grid ion thruster radiofrequency discharge.

13. Power supply assembly (1) for grid ion thruster (2), the thruster comprising: a plasma enclosure (20), a grid screen (24) which is located in front of or at an exit opening of the plasma enclosure, and an acceleration grid (25) which is parallel to the screen grid and located on one side of said screen grid opposite the plasma enclosure, the power supply assembly (1) comprising: - a common reference node (CRP); - a positive power supply unit (12), intended to be connected to the screen grid (24) of the thruster (2) to provide said screen grid, during operation of said thruster, a voltage positive electrical current (Vpuy) relative to the common reference node (CRP); - a negative power supply unit (13), intended to be connected to the acceleration grid (25) of the thruster (2) to provide said acceleration grid, during operation of said propellant, a negative electrical voltage (Vxuv) with respect to the common node of reference (CRP); - a variation module (15), arranged to vary the electrical voltage positive (Vpuy) supplied by the positive power supply unit 112): - measuring and recording modules (16), arranged to measure and record values ​​of an electric current flowing from the grid acceleration (25) to the power supply assembly (1), called acceleration grid current (Inuy); and - a controller (17), configured to activate the variation module (15) and the measuring and recording modules (16), so that variable values ​​are produced successively for the voltage positive electrical voltage (Vpuy) while negative electrical voltage (Vnwuv) is kept constant, and that a value of the acceleration grid current {(Ivnv) be measured and recorded for at least one provided value of the positive electrical voltage, when the power supply assembly (1) is connected to the thruster (2) to enable operation of said propellant, the power supply assembly (1) being adapted to execute a method which is in accordance with any one of claims | to 12.

14. A plasma propulsion system (10) comprising: - a grid ion thruster (2), the thruster comprising: a plasma enclosure (20), a screen grille (24) which is located in front of or at the level of an outlet opening of the plasma enclosure, and a grid acceleration (25) which is parallel to the screen grid and located on one side of said screen grid opposite the plasma enclosure, so that when from the operation of said propellant, positive ions which are produced in the plasma enclosure pass through the screen grid then the grid acceleration; and - a power supply assembly (1) which conforms to the resale- dication 13, and which is connected to the thruster (2) to produce a func- operation of said propellant.

15. Space vehicle (100), comprising a plasma propulsion system (10) which is in accordance with claim 14.