System comprising a space vehicle and method for determining an electrical phenomenon of a solar generator

Abstract

TITLE: System comprising a space vehicle and method for determining an electrical phenomenon of a solar generator The present invention relates to a system (1) comprising a space vehicle (3) with at least one solar generator (9), each solar generator comprising a plurality of sections and being suitable for supplying electric current to at least one electrical energy consumer (12, 14, 16, 18) of the space vehicle (3); at least one pair (24) of conductors, the pairs (24) of conductors comprising at least one conductor (24a, 24b) for each polarity; the space vehicle (3) comprising for each section (22) at least one sensor (28, 28a, 28b, 28c, 28d);and at least one device (16, 20) for detecting an electrical phenomenon, the device for detecting an electrical phenomenon being capable of: determining (1000) a common-mode current for each section (22) and / or the derivative of the common-mode current for each section (22), and determining the presence of an electrical phenomenon and the section(s) (22) of the solar generator (9) affected by the electrical phenomenon. Figure for abstract 2;

Description

Title of the invention: System comprising a space vehicle and method for determining an electrical phenomenon of a solar generator

[0001] The present invention relates to a system comprising a space vehicle with at least one solar generator. Furthermore, the present invention relates to a computer-implemented method for determining an electrical phenomenon of a solar generator.

[0002] Solar generators include solar cells with protective coverglass. In orbit, the coverglass becomes positively charged, and together with the solar cells, they form a capacitor that can discharge provided a plasma is present on its surface. This initial plasma bubble is generated by an electrostatic discharge (ESD) or following a micrometeorite or orbital debris (MMOD) impact. This plasma is then sustained by the discharge current from the coverglass, which is concentrated at the cathode spot that emits electrons. The discharge from the coverglass thus propagates from the cathode spot. The propagation of the electrostatic discharge from the coverglass onto a wing of a solar generator is called a flashover (FO). An electrostatic discharge can then create an electric arc.

[0003] US 9,190,836 B2 uses detection by variation of the current ratio between the nominal and redundant lines, with Hall effect probes or magnetic devices.

[0004] The article “Mapping of the Appleton Anomaly Using Arc Detectors on Starlink Group 6-1 Satellites” presented at the 17th Spacecraft Charging Technology Conference, Palais des Papes, Avignon, France, June 17-21, discloses a system for detecting an arc. A detector is positioned on the satellite body which does not allow localization of the section emitting the flashover, the discharges

[0005] A device for protecting solar panels is known in EPI 709 504 B1. The device includes a voltage drop detection circuit and an arc extinguishing circuit.

[0006] Other simulations have been carried out on the ground, for example in Japan with the Kyushu Institute of Technology, see Okumura, T. et al. “Flashover plasma characteristics on 5m² solar array panels in a simulated plasma environment of geostationary orbit and low Earth orbit,” AIAA 2010-1602, 48th Aerospace Science Meeting, Orlando, USA, January 2010, and in Europe, see Virginie Inguimbert et al. “Measurements of the Flashover Expansion on a Real-Solar Panel—Preliminary Results of EMAGS3 Project” » IEEE TRANSACTIONS ON PLASMA SCIENCE, VOL. 41, NO. 12, DECEMBER 2013.

[0007] But due to the limitation of surface area, in particular, of simulation tests carried out on the ground it is difficult to validate modelling tools because the architecture and the modalities are very different compared to actual use in flight.

[0008] The aim of the invention is to provide a system and a method for detecting electrical phenomena in an orbiting solar generator, for example, a flashover. In particular, the aim of the invention is to provide a system and a method for passivating an electric arc that can be operated autonomously or activated after detection of the electric arc.

[0009] To this end, the invention relates to a system comprising: a space vehicle with: at least one solar generator capable of generating an electric current, each solar generator being covered by a transparent protective material, each solar generator comprising a plurality of sections; at least one electrical energy consumer, each solar generator being capable of supplying electrical current to at least one electrical energy consumer of the space vehicle; at least one pair of electrical conductors associated with each section to supply the electrical current from each section of each solar generator to the respective electrical power consumers, the pairs of conductors comprising at least one conductor for each polarity; the space vehicle comprising, for each section associated with at least one pair of conductors, at least one sensor to measure a value dependent on the current(s) in the conductors; and the system comprising at least one device for detecting an electrical phenomenon of the solar generator, the device for detecting an electrical phenomenon being specific to: determine a common-mode current for each section and / or the derivative of the common-mode current for each section, and determine the presence of an electrical phenomenon and the section(s) of the solar generator concerned by the electrical phenomenon determined from the common mode current of each section and / or the derivative of the common mode current of each section.

[0010] According to other advantageous aspects of the invention, the system comprises one or more of the following features, taken individually or in all technically possible combinations: • each solar generator comprises at least two sections; • each solar generator forms a photovoltaic wing; • System according to any one of the preceding claims, characterized in that the space vehicle is a satellite and / or a space probe; • the sensor is a sensor measuring the derivative of a current, in particular a sensor including a toroidal coil, or a current sensor; • the system comprising exactly one sensor per pair of conductors; • the electrical phenomenon is a discharge propagation, a secondary arc, a primary arc without discharge propagation, and / or transient currents between solar generator and electric propulsion; • the space vehicle is capable of making the relevant section(s) determined passive, for example by short-circuiting the relevant section(s) determined; • the device for detecting an electrical phenomenon being capable of calculating the sum of all common-mode currents of each section and / or of calculating the sum of all derivatives of the common-mode currents of each section of a solar generator, and using the result of the calculation of the sum of all common-mode currents of each section and / or the sum of all derivatives of the common-mode currents of each section of a solar generator, the device for detecting an electrical phenomenon being capable of determining the type of electrical phenomenon among a detection of a discharge propagation, a secondary arc, a primary arc without discharge propagation, and / or transient currents between solar generator and electric propulsion; • The space vehicle includes a power conditioning unit and the power conditioning unit includes the sensors. • at least one device for detecting an electrical phenomenon of the solar generator is arranged on the ground; • at least one device for detecting an electrical phenomenon of the solar generator is installed in the spacecraft; and / or • a first part of a device for detecting an electrical phenomenon of the solar generator is arranged in the space vehicle and a second part of the device for detecting an electrical phenomenon of the solar generator is arranged on the ground.

[0011] The invention also relates to a method for determining an electrical phenomenon of a solar generator of a space vehicle with: at least one solar generator capable of generating an electric current, each solar generator being covered by a transparent protective material, each solar generator comprising a plurality of sections; at least one electrical energy consumer, each solar generator being capable of supplying electrical current to at least one electrical energy consumer of the space vehicle; at least one pair of electrical conductors associated with each section to supply the electrical power from each section of each solar generator to the respective electrical power consumers, the pairs of conductors comprising at least one conductor for each polarity; the process comprising: to obtain a measurement in each pair of conductors, a value depending on the current(s) in the conductor(s); and determine the common-mode current for each section and / or the derivative of the common-mode current for each section, and determine the presence of the electrical phenomenon and the section(s) of the solar generator concerned by the electrical phenomenon, from the common mode current of each section and / or the derivative of the common mode current of each section.

[0012] The invention also relates to a computer program comprising software instructions which, when executed by a computer, implement a method as defined above.

[0013] The invention will become clearer upon reading the following description, given solely by way of non-limiting example, and made with reference to the drawings in which:

[0014] [Fig-1] [Fig.1] schematically shows a system according to one embodiment;

[0015] [Fig.2] [Fig.2] schematically shows a solar generator;

[0016] [Fig.3] [Fig.3] schematically shows a sensor of a first mode of realization ;

[0017] [Fig.4] [Fig.4] schematically shows a sensor of a second mode of realization ;

[0018] [Fig.5] [Fig.5] schematically shows a sensor of another embodiment;

[0019] [Fig.6] [Fig.6] is a top view of the sensor in [Fig.5];

[0020] [Fig.7] [Fig.7] schematically shows a sensor of another mode of realization ;

[0021] [Fig.8] [Fig.8] shows a graph of the current and the current derivative of a section of the solar generator that emits a FO; and

[0022] [Fig.9] [Fig.9] shows a graph of the current and the derivative of the current of a section of the solar generator that collects current from a fiber optic cable; and

[0023] [Fig. 10] [Fig. 10] shows a flowchart of a process according to a mode of realization.

[0024] Figure 1 schematically shows a system 1 according to one embodiment. The system 1 comprises a space vehicle 3. The space vehicle 3 orbits the Earth 5.

[0025] In one embodiment the space vehicle 3 is a satellite.

[0026] The spacecraft 3 comprises a body 7. In addition, the spacecraft 3 comprises One or more solar generators (9). Each solar generator is connected to the body (7). Each solar generator is capable of generating an electric current. The term "solar generator" used below refers to a solar generator wing.

[0027] In one embodiment, the solar arrays 9 are respectively connected by a solar array drive mechanism 10 to the body 7 of the spacecraft 3. The drive mechanism 10 is designed to optimally align the solar arrays 9 with respect to the sun. For example, this mechanism includes a motor, and in particular a geared motor. Furthermore, the drive mechanism 10 includes a portion of the conductors that transmit the electric current generated by the respective solar arrays 9 to the body 7 of the spacecraft 3.

[0028] In another embodiment, the solar generator(s) 9 are fixed to the body 7 in a fixed manner relative to the body 7. In other words, the solar generator(s) 9 cannot be moved relative to the body 7.

[0029] The body of the space vehicle 3 includes at least one electrical power consumer 12.

[0030] For example, the electrical energy consumer 12 is an observation instrument and / or a means of telecommunications.

[0031] Each solar generator 9 being suitable for supplying electric current to at least one electrical energy consumer 12 of the space vehicle 3.

[0032] In addition, the body 7 of the spacecraft 3 includes a power conditioning unit (PCU) 14 and / or a power bus. The power conditioning unit 14 regulates the electrical power from the various electrical power sources, including solar generators 9 and / or batteries. In particular, the power conditioning unit 14 regulates the current and / or voltage on the power bus. The electrical power consumer(s) 12 are, for example, connected to the power bus.

[0033] The power bus is suitable for supplying current to or to each electrical energy consumer 12, for example instruments and / or means of telecommunications.

[0034] Optionally, the spacecraft 3 includes one or more devices 16 for detecting an electrical phenomenon of the solar generator. For example, in one embodiment, the spacecraft 3 includes, for each solar generator, 9 A device 16 for detecting a distinct electrical phenomenon. In another embodiment, a single device 16 for detecting an electrical phenomenon processes the electrical phenomena for all the solar generators 9 of the spacecraft 3. In this embodiment, the device 16 is a consumer of electrical energy. An electrical phenomenon is an electrical effect that is not related to the generation of an electric current by the solar generator 9. For example, the electrical phenomenon is a discharge propagation of the coverglasses on a solar generator (flashover), a secondary arc, a primary arc without discharge propagation, and / or transient currents between the solar generator and the electric propulsion system.

[0035] The electrical phenomenon detection device 16 is suitable for implementing a method for determining an electrical phenomenon of a solar generator of a space vehicle which will be described later.

[0036] The device 16 is an electronic circuit designed to manipulate and / or transform data represented by electronic or physical quantities in computer registers and / or memories into other similar data corresponding to physical data in register memories or other types of display devices, transmission devices or storage devices.

[0037] As specific examples, device 16 is implemented in the form of a A programmable logic component, such as an FPGA (Field Programmable Gate Array), or an integrated circuit, such as an ASIC (Application-Specific Integrated Circuit). In another example, device 16 is at least implemented through analog information processing, for example, a comparator-type detection circuit that can, for example, activate a circuit to passivate an electric arc.

[0038] Alternatively, when the method is implemented in the form of one or more software programs, i.e., in the form of a computer program, also called a computer program product, it is further capable of being stored on a computer-readable medium (not shown). The computer-readable medium is, for example, a medium capable of storing electronic instructions and being connected to a bus of a computer system. By way of example, the readable medium is an optical disc, a magneto-optical disc, ROM, RAM, any type of non-volatile memory (e.g., FLASH or NVRAM), or a magnetic card. A computer program comprising software instructions is then stored on the readable medium.

[0039] The spacecraft 3 is optionally equipped with a propulsion device 18. For example, the propulsion device 18 is an electric propulsion device, in particular a plasma propulsion unit (PPU). Propulsion Unit (in English). The propulsion unit 18 is, for example, one of the electrical power consumers of the spacecraft. In one embodiment, the propulsion unit 18 is connected to the power bus.

[0040] Furthermore, [Fig. 1] shows a ground-based data processing unit 20. The data processing unit is designed to receive data from the spacecraft 3. In particular, the data processing unit 20 is designed to store the data received from the spacecraft 3 in a database.

[0041] The data processing unit 20 is an electronic circuit designed to manipulate and / or transform data represented by electronic or physical quantities in computer registers and / or memories into other similar data corresponding to physical data in register memories or other types of display devices, transmission devices or storage devices.

[0042] As specific examples, the data processing unit 20 is implemented in the form of a programmable logic component, such as an FPGA (Field Programmable Gate Array), or an integrated circuit, such as an ASIC (Application-Specific Integrated Circuit), microprocessor, or microcontroller. In another example, device 20 is at least implemented through analog information processing.

[0043] Alternatively, when the method is implemented in the form of one or more software programs, i.e., in the form of a computer program, also called a computer program product, it is further capable of being stored on a computer-readable medium (not shown). The computer-readable medium is, for example, a medium capable of storing electronic instructions and being connected to a bus of a computer system. By way of example, the readable medium is an optical disc, a magneto-optical disc, ROM, RAM, any type of non-volatile memory (e.g., FLASH or NVRAM), or a magnetic card. A computer program comprising software instructions is then stored on the readable medium.

[0044] In one embodiment, the data processing unit 20 includes a device for detecting an electrical phenomenon of the solar generator. In this case, the data received from the spacecraft 3 by the data processing unit 20 are the data necessary to determine, at least in part, the electrical phenomenon of the solar generator.

[0045] Fig. 2 schematically shows a solar generator 9.

[0046] Each solar generator 9, in particular the photovoltaic cells, is covered by a transparent protective material, for example a protective glass or protective window. The protective glass or protective window, called a "cover The term "glass" in English refers to a material suitable for use in space. For example, protective glass is designed to withstand radiation in orbit. Protective glass is typically thin. In one embodiment, the thickness of the protective glass is between 75 micrometers and 1 millimeter. The transparent protective material is suitable for covering one or more photovoltaic cells.

[0047] Each solar generator 9 comprises a plurality of sections 22. Each section comprises one or more strings. Each solar generator comprises at least 2 sections, for example at least 3, 4, or 5 sections, in particular between 2 and 90 sections.

[0048] In one embodiment, one or more blocking diodes are mounted on each string. The blocking diodes block reverse current in the string in question, for example, in the case of a shaded string, a string insulation fault, or in the event of an EPS (Electrical Power Subsystem) failure, where current would otherwise flow from the bus to the section. For example, the diodes isolate one string from the other strings in the same section, for example, in the event of a string failure.

[0049] The solar generator(s) 9 are suitable for supplying electric current to the body 7 of the space vehicle 3, in particular to the power conditioning unit 12 (PCU) and / or the power bus.

[0050] At least one pair 24 of electrical conductors is associated with each section 22 of each solar generator to supply the electric current from each section 22 to the respective electrical power consumer(s) 12, 14, 16, 18.

[0051] Each pair 24 of conductors comprises at least one conductor of a first polarity 24a, for example "+", and at least one conductor of a second polarity 24b, for example "-".

[0052] According to one embodiment, for each polarity the pair 24 of conductors comprises several redundant conductors. The redundant conductors are, for example, separated from each other at least in certain places.

[0053] The second polarity conductors 24b are in an example connected to a ground reference point 26, for example the 0V satellite.

[0054] The conductors of the first polarity 24a are for example respectively connected to the power conditioning unit 14, to a power supply bus and / or to an electrical power consumer 12.

[0055] The space vehicle 3 includes for each pair 24 of conductors at least one sensor 28 to measure a value dependent on the current(s) in the conductors 24a, 24b.

[0056] The sensor(s) 28 are outside the power conditioning unit 14.

[0057] According to one embodiment, the sensor(s) 28 are integrated into the power conditioning unit 14.

[0058] In one embodiment, the sensor 28 is a sensor measuring the derivative of a current, in particular a sensor comprising a toroidal coil. Therefore, the value depending on the current(s) in the conductors 24a, 24b can be a derivative of the current. The voltage measured across the toroidal coil is a representation of the derivative of the current flowing through the sensor 28.

[0059] Figure 3 schematically shows a sensor 28a of a first embodiment. The sensor 28a is suitable for determining a value dependent on the current(s) in the conductors 24a, 24b.

[0060] The sensor 28a comprises a ferromagnetic core 30a. The ferromagnetic core can have a variable geometry.

[0061] In [Fig. 3], the ferromagnetic core 30a has a circular annular shape or a torus shape. Other shapes are also possible, for example, the shapes described later, an "El" shape or an "EE" shape.

[0062] A coil 32a is mounted on at least a part of the ferromagnetic core 30a. At least one conductor of the first polarity 24a and at least one conductor of the second polarity 24b pass through the ferromagnetic core 30a, in particular such that the current directions in the conductors 24a and 24b are opposite. In one example, all the conductors of the first polarity 24a and all the conductors of the second polarity 24b pass through the ferromagnetic core 30a.

[0063] According to one embodiment, the sensor 28a includes an insulating barrier between the conductor(s) of the first polarity 24a and the conductor(s) of the second polarity 24b, in particular to electrically isolate the conductors 24a and 24b from each other.

[0064] The sensor 24a measures the derivative of the difference in currents (dl / dt) flowing in conductors 24a and 24b, in particular through coil 32a. In other words, the sensor 24a measures the derivative of the vector sum of the currents flowing in conductors 24a and 24b.

[0065] The vector sum of the current in each pair of conductors is also called the common-mode current or zero-sequence current. In an example, when the electrical phenomenon is detected, it is a transient common-mode current.

[0066] This measurement method allows segregation between the power circuits of the solar generator and the current measurement circuits, in particular the sensors 28. Therefore, the reliability of the mission is not compromised where a failure of either the sensor or the power supply system in the form of the solar generator 9 and / or the conductors 24a, 24b propagates to other systems.

[0067] In one embodiment, the sensor 28a has a ferromagnetic core that can be opened during assembly to insert the conductor(s) 24a, 24b. Then the ferromagnetic core is closed.

[0068] According to one embodiment, two sensors 28a are used for each section 22. The first sensor 28a measures a value dependent on the current(s) in the conductor(s) of the first polarity 24a, and the second sensor 28a measures a value dependent on the current(s) in the conductor(s) of the second polarity 24b. In this case, the first sensor 28a measures the derivative of the current flowing in the conductor(s) 24a, and the second sensor 28a measures the derivative of the current flowing in the conductor(s) 24b. For example, in this case, the coils 32a of the first and second sensors are connected in series so as to obtain the derivative of the difference in currents (dl / dt) flowing in the conductors 24a and 24b. In another embodiment, the series connection is achieved through data processing, for example, in the electrical phenomenon detection device 16.

[0069] Figure 4 schematically shows a sensor 28b of another embodiment. Sensor 28b is suitable for determining a value dependent on the current(s) in conductors 24a, 24b. In sensor 28b, the characteristics of sensor 28b having the same functions have the same reference numerals as the embodiment of sensor 28a in Figure 3, except with a "b" following the numeral instead of an "a" (except for conductors 24a, 24b, which will have the same characteristics as in Figure 3). Unlike Figure 3, sensor 28b is provided with a ferromagnetic core 30b in the shape of a square ring.

[0070] In other embodiments, the shape of the ferromagnetic core may have yet other shapes, for example, a hexagonal, octagonal or oval shape.

[0071] Figure 5 schematically shows a sensor 28c of another embodiment in an axial view. The axis corresponds to the axis of conductors 24a, 24b. Figure 6 is a top view of the sensor 28c of Figure 5. In the sensor 28c, the characteristics of the sensor 28c having the same functions have the same reference numerals as the embodiment of the sensor 28a of Figure 3, except with a "c" following the numeral instead of an "a" (except for conductors 24a, 24b, which will have the same characteristics as in Figure 3). The sensor 28c comprises a double-toroidal ferromagnetic core 30c and a single coil 32c. In other words, the ferromagnetic core 30c comprises a first torus and a second torus. The first and second torus have a common cross-section 36c. Each torus has a circular shape.The first torus is traversed by at least one conductor of the first polarity 24a and the second torus is traversed by at least one conductor of the second polarity 24b, in particular in such a way that the direction of the current in the conductor(s) of the first polarity 24 and the direction of the current in the conductor(s). conductors of the second polarity 24b are approximately parallel to each other, in particular as illustrated in [Fig.6].

[0072] Figure 7 schematically shows a sensor 28d of another embodiment. The sensor 28d is suitable for determining a value dependent on the current(s) in the conductors 24a, 24b. In the sensor 28d, the characteristics of the sensor 28c, which has the same functions, have the same reference figures as the embodiment of the sensor 28c in Figure 5, except with a "d" following the number instead of a "c" (except for the conductors 24a, 24b, which will have the same characteristics as in Figure 3). Unlike Figure 3, the sensor 28d is provided with a ferromagnetic core 30d, where each torus has the shape of a square ring.

[0073] According to one embodiment, all the conductors 24a, 24b associated with each section 22 pass through the sensor 28a, 28b, 28c, 28d, in particular the ferromagnetic core 30a, 30b, 30c, 30d.

[0074] In one embodiment, the sensor(s) is / are arranged electrically between the drive mechanism 10 (SADM) and the power conditioning unit 14 (PCU). For example, the sensor(s) is / are arranged on an electronic board in the sensing device 16 which incorporates a ferromagnetic core.

[0075] In another embodiment, the sensor is a current sensor. For example, each current sensor is specific to measuring the current value. The sensors are, for example, integrated into the power conditioning unit 14. For example, the sensor may be provided with a resistor whose potential difference is measured across its terminals and compared to a reference, for example, an inductor coupled across whose terminals the voltage variation induced by a current variation is measured.

[0076] According to one embodiment, the sensors are located in the body 7 of the spacecraft between the drive mechanism 10 (SADM) and the power conditioning unit 14 (PCU). A first current sensor is used, for example, to measure the current in the first-polarity conductor 24a, and a second current sensor is used, for example, to measure the current in the second-polarity conductor 24b. Then, the vector sum of the currents is calculated, for example, by the electrical phenomenon detection device 16.

[0077] According to the invention, measurements taken by currents in conductors 24a, 24b and sensors 28, 28a, 28b, 28c, 28d do not disrupt the operation of the solar generator 9, nor do they degrade or compromise the reliability of the spacecraft mission. Furthermore, the invention provides protection for conductors 24, 24a, 24b between the blocking diode of the solar generator 9 and the PCU 14, including the SADM 10, against arcing.

[0078] According to one embodiment, the current is considered positive if it flows in the conductors 24a, 24b from the solar generator to the body of the satellite.

[0079] During normal operation, i.e. if there is no electrical phenomenon or FO, at any instant for a solar generator 9, the sum over all sections 22 of the vector sums of the currents of each pair of conductors is zero.

[0080] For example, in Figure 2, 11+12+13+14+15+16=0, with 11, 12, 13, 14, 15 and 16 being the vector sums of the currents in the sections of the solar generator flowing respectively in conductors 24a, 24b. More generally: _ n, with ^=12 u N the number of sections, pi being the vector sum of the currents in the pair of conductors 24 of section i.

[0081] Furthermore, during normal operation, the vector sum of the current in each pair of conductors is equal to zero. In the example of [Fig. 2]:

[0082] 11=12=13=14=15=16=0

[0083] More generally = 0, for all sections i.

[0084] Furthermore, during normal operation, each section 22 of the solar generator 9 generates a direct current. Therefore, the derivative of the vector sum of the current in each pair of conductors is zero. For example, in the example of [Fig. 2]

[0085] dll / dt = dI2 / dt = dI3 / dt = dI4 / dt = dI5 / dt = dI6 / dt= 0

[0086] More generally, _ n Qa is the derivative of the vector sum of the current I of dt -u each pair of conductors) for all sections i).

[0087] In the following, the current behavior for each section during an electrostatic discharge (Flashover - FO) is explained. FO is an electrical phenomenon.

[0088] During an electrostatic discharge propagation (Flashover - FO), there is a section 22 of the solar generator 9 which emits the FO and one or more sections which collect(s) the current from the FO.

[0089] During a FO, the vector sum of the currents in the conductor pair of the section emitting the FO, particularly with a cathode spot on that section, is positive, and for the sections collecting the FO current, the vector sum of the currents in the conductor pairs is negative. Therefore, during a FO, there are certain sections 22 where the vector sum of the currents in the conductor pairs is not zero.

[0090] Fig. 8 shows a graph of the vector sum of currents and the derivative of the vector sum of currents of a section 22 which emits an FO from the solar generator 9.

[0091] Figure 9 shows a graph of the vector sum of the currents and the derivative of the vector sum of the currents of a section 22 of the solar generator 9 which collects the current from a FO.

[0092] Before and during a FO, the sum of the vector sums of the currents in each section is zero. In the example in [Fig. 2]: 11+12+13+14+15+16= 0. More generally:

[0093] 2NI2=0(D-

[0094] with N the number of sections, pi the vector sum of the currents in the pairs of conductors 24 of section i.

[0095] Moreover, before and during a FO, the sum of each derivative of the vector sum of the currents of each pair of conductors 24 of a section 22 of a solar generator 9 is equal to zero. In other words, in the example of [Fig. 2]: d1 / dt + dI2 / dt + dI3 / dt + dI4 / dt + dI5 / dt + dI6 / dt = 0. More generally:

[0096] dl^ _ (2).

[0097] with N the number of sections, pi the vector sum of the currents in the conducting pairs 24 of section i and t the time.

[0098] Equations (1) and (2) apply unless there is yet another electrical phenomenon or another possible electrical circuit, for example a leak in a conductor or if a current is collected by the electric propulsion 18.

[0099] In Figures 8 and 9 the start of the FO is marked by A, the time of the maximum (absolute) of the vector sum of the current is marked respectively by B and B', and the extinction time of the FO is marked by C.

[0100] During the start-up of the FO, the derivative of the vector sum of the current of the pair of conductors 24 of the section 22 which emits an FO has a positive peak (at time A on the [Fig.8]), and during the extinction time of the FO the derivative of the vector sum of the current of the pair of conductors 24 of the section 22 which emits an FO has a negative peak ([Fig.8]).

[0101] Estimating the maximum of the vector sum of the FO current is possible by calculating an integral between A and B, which corresponds to the area following the FO dl / dt curve between A and B. When there is an FO, there is only one section 22 that emits and the other section(s) 22 collect the current.

[0102] During the start-up of the FO, the derivative of the vector sum of the current of the pair of conductors 24 of each section 22 which collects an FO has a negative peak (marked by A in [Fig.9]), and during the extinction time of the FO the derivative of the vector sum of the current of the pair of conductors 24 of the section 22 which collects an FO has a positive peak (marked by C on [Fig.9]).

[0103] At point B', the derivative of the vector sum of the current in each pair of conductors 24 of a section 22 that collects a FO is zero. It should be noted that the further the collecting sections 22 are from the section that emits the FO, the more the The current peak time (time B' in [Fig. 9]) is large. This corresponds to the propagation time of the optical flow between the cathode spot of the emitting section and the various collecting sections. The current peak time is also called the peak. Furthermore, the further the collecting sections are from the emitting section, the lower the dl / dt (point A in [Fig. 9]) at startup.

[0104] The determination of the presence of an electrical phenomenon is described below. Figure 10 shows a flowchart of a process according to one embodiment.

[0105] In a first step 1000, a common-mode current for each section 22 associated with at least one pair of conductors 24a, 24b and / or the derivative of the common-mode current for each section 22 associated with at least one pair of conductors 24a, 24b are determined. In one embodiment, the common-mode current common for each section 22 and / or the derivative of the common mode current for each section 22 is / are recorded.

[0106] The vector sum of the current in each section corresponds to a common-mode current and can also be called a zero-sequence current.

[0107] For example, before a discharge propagation detection (FO), during normal operation, the sum of the currents in each section 22 is constant. More generally, the derivative of the common-mode current I in each section 22 is constant. for all sections i.

[0108] In step 1010, it is determined whether the common-mode current I of each section is above a threshold s(I; > s > 0), or whether the derivative of the common-mode current I of each section is above a threshold d(dQ > rf > (p' ^ans ce cas' 'cs sections that emit are detected.

[0109] In another embodiment, all the sections that collect are detected. In this case, the currents and thresholds are negative, for example (I; < sc <

[0110] If > d>Q For at least a section 22, with d being a threshold, or if I; > s > 0 For at least one section 22, where s is a threshold, step 1020 determines whether other conditions for a FO are met, specifically whether all conditions are met. In the embodiment where the collecting sections are detected by the comparison indicated above, it is also determined whether other conditions for a FO are met.

[0111] The conditions are as follows:

[0112] A) if at the time of the start of the FO (point A on figures 8 and 9): a) x N dlj _ up to noise and measurement uncertainties (equation (2)); A=1 dt - u b) dI / dt>0 for a section (which emits); c) dl / dt <0 for at least one other section (which collects); d) dI / dt=0 for the other sections; and e) the sections that emit and collect are all adjacent.

[0113] Regarding condition e), for example, if sections 2, 3 and 4 are affected by conditions b) and c), condition e) is met. If sections 1, 3, 4 are affected by conditions b) and c), condition e) is not met.

[0114] B) If from the moment of start-up (moment A in Figures 8 and 9) until extinction (moment C in Figures 8 and 9) for any instant t: a) x _ n within noise and measurement uncertainties; dt “ u b) v AT j _ n, within the limits of noise and measurement uncertainties, the sum of the currents I(t), A=1 1 if they are not directly measured are calculated by integrating from tO (start-up time A in figures 8 and 9) to t; c) I>0 for a section (which emits);

[0115] d) I <0 for at least one other section (which collects); e) I = 0 for the other sections; and f) the sections that emit and collect are all adjacent.

[0116] Regarding condition e), for example, if sections 2, 3 and 4 are affected by conditions b) and c), condition e) is met. If sections 1, 3, 4 are affected by conditions b) and c), condition e) is not met.

[0117] C) If at the time of FO extinction (time C in figures 8 and 9): a) an extinction of synchronized dl / dt (the C times are at the same time for all sections that emit and collect)

[0118] b) dI / dt<0 for one section (which emits);

[0119] c) dl / dt > 0 for at least one other section (which collects); and d) dVdt=0 for the other sections.

[0120] D) If a. The further the collecting sections are from the emitting section, the larger the peak (point B' on [Fig.9]), b. Optional condition: The further the collecting sections are from the emitting section, the lower their dl / dt at startup.

[0121] In one embodiment, particularly if the space vehicle is equipped with an electric propulsion device 18, during the start-up or shutdown phase of the propulsion device 18, formulas (1) and (2) (see conditions A) a), B) a), and B) c)) are modified with the following equations:

[0122] T - T (rXet ^=1^'” ~lc

[0123] yN dl1 =

[0124] With being the collected current. The collected current between a section of the solar generator 9 and the propulsion device. The collected current can be determined during the start-up or shutdown phases without FO. In this case, dlc / dt < 0 at propulsion start-up and dlc / dt > 0 at propulsion shutdown. For sections 22 that switch, i.e., that alternately switch between a short-circuited and unshort-circuited state, the signal dlc / dt depends on the switching frequency. The current variation related to the start-up of the propulsion device 18 is mainly provided by the capacitance of the power supply bus CBus (equivalent bus capacitance).At the solar generator level this translates into the connection, for example with an electronic switch, of a section on the power bus in order to provide the operating current to the electrical energy consumers and to provide the current necessary to recharge the capacitance of the CBus bus.

[0125] In an optional step 1030, in the event of the detected electrical phenomenon, here the FO, the system, in particular the electrical phenomenon detection system 16 and / or the data processing unit 20, determines one or more of the following information: a date of the electrical phenomenon, the solar generator concerned by the respective electrical phenomenon, the section 22 which emits the FO, the section or sections 22 which collect the FO, the maximum current of the FO of the emitting section, the duration of the FO, in particular with the current of the emitting section 22, and / or the time-dependent shape of the FO current with the current of the emitting section.

[0126] All or part of this processing can be carried out on the space vehicle and completed on the ground, for example in the data processing unit 20.

[0127] In one embodiment, the information is supplemented with, at the time of the electrical phenomenon: the electrical state of the sections of the solar generator (on, off, for example short-circuited, switching), solar generator in eclipse, solar generator illuminated, shading of sections 22, state of the electric propulsion 18, current of one or more sections, and transients of the electric propulsion.

[0128] This additional processing can be done on board the space vehicle and / or on the ground, in particular in the data processing unit 20.

[0129] This information will allow us to have a distribution of the characteristics of the observed FOs, which will be useful in defining the FO to be simulated during ground qualification tests.

[0130] This information may also be useful for correlating events observed in flight with electrical phenomena determined on this space vehicle.

[0131] This information can be used to improve theoretical models.

[0132] In an optional step 1040, in the event of a detected electrical phenomenon, particularly a detected FO, the electrical phenomenon detection device 16 is designed to short-circuit the section 22 emitting the FO and / or adjacent sections 22 of the emitting FO, for example, via the power conditioning unit 14. The number of adjacent sections 22 depends on their distance from the emitting FO section 22. This prevents an associated secondary arc and degradation of the solar generator 9. The duration of the short-circuit of the sections 22 depends on the duration of the detected FO and / or the size of the solar generator. For example, the short-circuit duration is at least 10 ms, for example, at least 20 ms.

[0133] According to one embodiment, after the short-circuit duration has expired, the short-circuited sections 22 are reconnected to the power bus by the power conditioning unit. Therefore, the electrical phenomenon detection device 16 is an arc-extinguishing device that activates upon detection of a short-circuit current (SOC). This arc-extinguishing device has the advantage of reducing the duration of secondary arcs in orbit, even if the SOC continues its propagation independently, with a duration longer than that of the secondary arcs. This reduces the associated constraints for qualification tests for large-area solar generators.

[0134] The invention makes it possible to increase the differential voltage between solar cells above 30V and to increase the maximum current of a string above 1.5A.

[0135] According to one embodiment, another electrical phenomenon, in particular a (secondary) arc without a FO, is detected, for example, one triggered by the impact of a micrometeoroid and orbital debris (MMOD). Following this detection of a secondary arc, the section(s) concerned are short-circuited to extinguish the arc, as described above.

[0136] In the case of a secondary arc between cells of different cross-sections 22, the common-mode current I of the pairs of conductors of the cross-sections concerned is not zero and the sum of the cross-sections is equal to zero, i.e. j _ q. This detection does not does not function during the start-up or shutdown phase of the propulsion device 18.

[0137] Therefore, the electrical phenomenon detection device 16 is suitable for calculating the sum of the common-mode currents of the pairs of conductors 24, 24a, 24b. If S~x=0, the common-mode currents of the two sections between which the secondary arc is established are non-zero, one of the sections having a positive common-mode current and the other section having a corresponding negative common-mode current, and the common-mode currents of the other sections being zero, the detection device of an electrical phenomenon 16 is suitable for determining that there is an arc between two sections without FO.

[0138] In the case of a secondary arc between cells of the same section 22, the common-mode current of each section is zero, so j = q, and a variation in common-mode current is detected, in particular by the power conditioning unit 14. In this case, the power conditioning unit informs the detection device 16. The detected variation corresponds to a partial or total loss of a string of the section concerned, for example, if a blocking diode is used per string. If these conditions are met, the detection device 16 determines that there is an arc between cells of the same section without FO.

[0139] In another embodiment, arcs located between the blocking diodes and the power conditioning unit 14, for example at the drive mechanism (SADM), are detectable with a drop in current and / or voltage on the section concerned, for example by the power conditioning unit 14 or by the electrical phenomenon detection device 16.

[0140] In another embodiment, another electrical phenomenon, particularly when a primary arc without a FO (Fuel Off) is detected, for example a blow-off. This is a discharge of the spacecraft's capacitance into space, which can pass through the conductor of the solar generator 9. In this case, it is a discharge of a rapid negative charge from the satellite 3 from a site on the satellite that is not necessarily on a solar generator 9. The solar generators contribute to this overall discharge, so a discharge current occurs between the sections and the body 7 of the satellite 3. In this case, the electrical phenomenon detection device determines whether h < 0 for all sections (current to all sections of the solar generator). If this is the case, the detection device 16 determines that there is a primary arc without a FO, the site of which is not on a solar generator.

[0141] List of reference signs

[0142] 1 system

[0143] 3 satellite

[0144] 5 Earth

[0145] 7 bodies

[0146] 9 solar generator

[0147] 10 drive mechanism

[0148] 12 electrical energy consumer

[0149] 14 the energy conditioning unit

[0150] 16 Device for detecting an electrical phenomenon of the solar generator

[0151] 18 propulsion device

[0152]

[0153]

[0154]

[0155]

[0156]

[0157]

[0158]

[0159]

[0160]

[0161] 20 data processing unit 22 section 24 pair of conductors 24a, 24b conductor 26 ground 28, 28a, 28b, 28c, 28d sensor 30a, 30b, 30c, 30d ferromagnetic core 32a, 32b, 32c, 32d coil 34a, 34b insulation barrier 36c, 36d common section

Claims

1.

2.

3. Demands System (1) comprising: a space vehicle (3) with: at least one solar generator (9) capable of generating an electric current, each solar generator being covered by a transparent protective material, each solar generator comprising a plurality of sections; at least one electrical energy consumer (12), each solar generator being capable of supplying electrical current to at least one electrical energy consumer (12, 14, 16, 18) of the space vehicle (3); at least one pair (24) of electrical conductors associated with each section to supply the electric current from each section of each solar generator to the respective electrical power consumers, the pairs (24) of conductors comprising at least one conductor (24a, 24b) for each polarity; characterized in that the spacecraft (3) comprises, for each section (22) associated with at least one pair (24) of conductor, at least one sensor (28, 28a, 28b, 28c, 28d) for measuring a value dependent on the current(s) in the conductors; and in that the system comprises at least one device (16, 20) for detecting an electrical phenomenon of the solar generator (9), the device for detecting an electrical phenomenon being suitable for: determining (1000) a common-mode current for each section (22) and / or the derivative of the common-mode current for each section (22), and determine the presence of an electrical phenomenon and the section(s) (22) of the solar generator (9) concerned by the electrical phenomenon determined from the common mode current of each section (22) and / or the derivative of the common mode current of each section (22). A system according to any one of the preceding claims, characterized in that each solar generator (9) comprises at least two sections. A system according to any one of the preceding claims, characterized in that each solar generator (9) forms a photovoltaic wing.

4. System according to any one of the preceding claims, characterized in that the space vehicle (3) is a satellite and / or a space probe.

5. System according to any one of the preceding claims, characterized in that the sensor (28, 28a, 28b, 28c, 28d) is a sensor measuring the derivative of a current, in particular a sensor comprising a toroidal coil, or a current sensor.

6. System according to any one of the preceding claims, comprising exactly one sensor (28, 28a, 28b, 28c, 28d) per pair of conductors (24a, 24b).

7. System according to any one of the preceding claims characterized in that the electrical phenomenon is a discharge propagation, a secondary arc, a primary arc without discharge propagation, and / or transient currents between solar generator and electric propulsion.

8. System according to any one of the preceding claims, characterized in that the space vehicle (3) is suitable for passivating the determined relevant section(s), for example by short-circuiting the determined relevant section(s) (22).

9. A system according to any one of the preceding claims, characterized in that the electrical phenomenon detection device (16) is suitable for calculating the sum of all common-mode currents of each section (22) and / or for calculating the sum of all derivatives of the common-mode currents of each section (22) of a solar generator, and using the result of the calculation of the sum of all common-mode currents of each section (22) and / or the sum of all derivatives of the common-mode currents of each section (22) of a solar generator, the electrical phenomenon detection device (16) is suitable for determining the type of electrical phenomenon among the detection of a discharge propagation, a secondary arc, a primary arc without discharge propagation, and / or transient currents between the solar generator and electric propulsion.

10. System according to any one of the preceding claims, characterized in that the space vehicle (3) comprises a power conditioning unit (14) and in that the power conditioning unit comprises the sensors (28, 28a, 28b, 28c, 28d).

11. System according to any one of the preceding claims characterized in that at least one device for detecting an electrical phenomenon of the solar generator (20) is arranged on the ground.

12. System according to any one of the preceding claims 1 to 10, characterized in that at least one device for detecting an electrical phenomenon of the solar generator (14) is arranged in the space vehicle.

13. System according to any one of the preceding claims 1 to 10, characterized in that a first part of a device for detecting an electrical phenomenon of the solar generator (14) is arranged in the space vehicle and a second part of the device for detecting an electrical phenomenon of the solar generator (14) is arranged on the ground.

14. A method for determining an electrical phenomenon of a solar generator of a space vehicle having: at least one solar generator (9) capable of generating an electric current, each solar generator (9) being covered by a transparent protective material, each solar generator comprising a plurality of sections (22); at least one electrical power consumer (12, 14, 16), each solar generator being capable of supplying the electric current to at least one electrical power consumer of the space vehicle; at least one pair of electrical conductors (24, 24a, 24b) associated with each section to supply the electrical power of each section of each solar generator to the respective electrical power consumers, the pairs of conductors comprising at least one conductor for each polarity;the method comprising: obtaining a measurement in each pair of conductors a value depending on the current(s) in the conductor(s) (24a, 24b); and determining the common-mode current for each section (22) and / or the derivative of the common-mode current for each section (22), and determining the presence of the electrical phenomenon and the section(s) of the solar generator concerned by the electrical phenomenon, from the common-mode current of each section (22) and / or the derivative of the common-mode current of each section (22).

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