Timer device for a spacecraft passivation system

Abstract

The invention relates to a timer device for activating a spacecraft passivation system. The timer device (2) is made up of a control device (4) coupled to a dosimeter (6). The dosimeter is capable of characterizing a dose of radiation to which a spacecraft is exposed during its mission, so that the control device (4) can know the dose of radiation absorbed by said dosimeter (6). The control device (4) is capable of activating a system for passivating a spacecraft when the dose of radiation received by said dosimeter (6) exceeds a reference value, corresponding to a predetermined time of flight on a predetermined orbit.

Description

Description Title of the invention: Timing device for a passivation system of a spacecraft Technical field to which the invention relates

[0001] The invention relates to the technical field of aerospace and, more particularly, to the passivation of end-of-life spacecraft. Technological background

[0002] Near-Earth space is cluttered with end-of-life artificial satellites that are likely to remain in orbit for several more decades. To avoid saturating near-Earth space, space agencies are limiting the time that satellites that are no longer operational can remain in low orbit to 25 years.

[0003] However, the explosion of non-operational satellites has been observed on several occasions, due to a malfunction of their electrical accumulators.

[0004] A satellite explosion generates a significant amount of small debris, the speed and direction of which are uncontrolled. Each piece of debris then presents a risk of collision with an operational spacecraft, a collision that could permanently damage it.

[0005] To avoid this risk of explosion, space agencies recommend passivating the satellites' electrical accumulators at the end of a mission. To achieve this, satellites now incorporate a passivation system, programmed to permanently deplete all electrical reserves on board the satellite, after a period memorized by the passivation system.

[0006] In order for the passivation system to carry out its mission autonomously, more precisely without instructions from outside the satellite, the passivation system is coupled with a timing device present in the satellite. Activation of the passivation system at the right time therefore depends on the accuracy of the timing device.

[0007] To preserve the integrity of the timing device during the satellite's launch into orbit, it is known to reinforce the interconnections between the various components of the timing device. These special arrangements generate additional costs and a manufacturing time for the timing device that are not negligible.

[0008] Timing devices are known to use electronic or atomic clocks to measure time. These clocks measure the passage of time based on the frequency of an electrical pulse generated by an electronic oscillator. However, the frequency of this electrical pulse varies depending on the temperature of the electronic oscillator. It is therefore essential to stabilize the temperature of the clock's electronic oscillator to enable regular time measurement. This is why digital or atomic clocks on board satellites are surrounded by a relatively thick thermal insulator to protect them from the frequent and extreme temperature variations to which satellites are exposed during their orbit around the Earth. The presence of this thermal insulator has the disadvantage of significantly increasing the size of electronic clocks in satellites.

[0009] Another drawback is that the operation of an electronic oscillator is likely to be impaired when exposed to electromagnetic radiation or energetic subatomic particles. In other words, when an electronic oscillator is exposed to electromagnetic radiation present in space, its operation deteriorates in an unpredictable and irreversible way over time. Electronic clocks integrated into orbiting satellites and operating using an electronic oscillator therefore present a risk of statistical time measurement error.

[0010] In order to limit this risk of statistical error, it is known to harden the electronic oscillator as well as the other components of the electronic clocks on board satellites. By "hardening", we mean that the electronic circuits are doubled or even tripled and / or that their dimensions are increased. This solution nevertheless has the disadvantage of significantly increasing the number of sensitive components of the timing device, components which are generally fragile and expensive. According to another disadvantage, the redundancy electronic circuits also significantly increase the size, weight and especially the power consumption of electronic clocks.

[0011] An alternative or complementary solution may then consist of placing the timing device in a housing to attenuate the radiation. However, this solution is not ideal because it increases the size of the timing device and complicates its connection to the passivation system.

[0012] For the reasons mentioned above, the electronic clocks currently used to enable the activation of a satellite passivation system at the end of its life are not suitable for small satellites or nanosatellites.

[0013] The invention aims to remedy this technical problem by proposing a new type of timing device, safer, less expensive and more compact, making it possible to activate a passivation system of a small satellite or nanosatellite at the end of the mission. Purpose of the invention

[0014] For this purpose, the invention proposes a timing device enabling the activation of a passivation system of a spacecraft. The term "timing device" means the ability of said device to activate a passivation system after a predetermined time.

[0015] Remarkably, the timing device comprises a control device coupled to a dosimeter. According to the invention, the dosimeter is capable of characterizing a dose of radiation to which a spacecraft is exposed during its mission. The radiation absorbed by the dosimeter is ionizing and / or non-ionizing.

[0016] The term "coupled" means that the control device is capable of taking note, over time, of the dose of radiation absorbed by the dosimeter, which can only increase over time, so that the control device can activate the passivation system of a spacecraft when the dose of radiation received by said dosimeter exceeds a reference value.

[0017] In other words, the timer device assesses the time spent in space by a spacecraft, based on the radiation dose received by a dosimeter.

[0018] Thus, in space, the timing device is safer to use than the electronic clocks described above, because the time is characterized in relation to the radiation present in the environment of the dosimeter and not in relation to an electronic oscillator whose operation is liable to be disturbed by the same radiation.

[0019] It should be noted that the dose of radiation to which the dosimeter is exposed, when the dosimeter moves in a defined orbit and duration, can be determined precisely or relatively precisely by a person skilled in the art. Therefore, the use of a dosimeter to characterize the flow of time in space is advantageously carried out in a deterministic manner and is not subject to reliability statistics as is the case for electronic clocks.

[0020] Thus, the timing device according to the invention may include fewer hardening and thermal insulation precautions, in comparison with the electronic clocks described above. For these reasons, the timing device according to the invention differs from the state of the art in that it is more economical, more compact and lighter.

[0021] According to another embodiment of the invention, the control device comprises a comparison unit connected to the dosimeter and a reference unit connected to the comparison unit. The reference unit is capable of communicating a reference value to the comparison unit, and the comparison unit is capable of activating a passivation system connected to the control device according to the invention.

[0022] The reference value corresponds to an estimated value of the radiation absorbed by the dosimeter, at the end of the mission of a spacecraft comprising the timing device according to the invention. As explained below, the reference value is easily estimable by a person skilled in the art, from the orbit traveled by the spacecraft during its mission and the absorption properties of the dosimeter and, possibly, the materials surrounding the dosimeter.

[0023] In order to increase the robustness of the timing device, all the elements making up the timing device are of the analog type and connected to each other by means of conductive tracks and / or electrically conductive wires.

[0024] For example, the reference unit may comprise one or more electrical resistors connected in series, powered by a voltage source, so as to deliver to the comparison unit an electrical voltage corresponding to the reference value.

[0025] According to a variant, the reference unit comprises one or more controlled switches, connected in parallel with one or more electrical resistors, so as to allow modification of the electrical voltage delivered by the reference unit to the comparison unit. By way of non-limiting example, the reference unit may be a voltage divider bridge comprising one or more switches controlled by the communication unit.

[0026] According to another embodiment of the invention, the control device comprises a communication unit capable of receiving a new reference value, and capable of substituting the reference value recorded by the reference unit with the new reference value.

[0027] According to another embodiment of the invention, the communication unit is capable of receiving a new reference value when the timing device is present in a spacecraft. Thus, advantageously, the passivation of a spacecraft equipped with the invention can be advanced or delayed depending on the circumstances of the mission.

[0028] Preferably, the new reference value is transmitted to the communication unit of the timing device, via radio frequency signals.

[0029] According to another embodiment of the invention, the dosimeter comprises a sensitive cell, composed of a material whose physical properties change proportionally to the dose of radiation absorbed.

[0030] According to another embodiment of the invention, the material composing the sensitive cell is capable of absorbing a dose of radiation whose value is between 1 krad and 1 Mrad, preferably between 1 krad and 10 krad.

[0031] According to another embodiment of the invention, the electrical conductivity of the material making up the sensitive cell evolves proportionally to the dose of radiation absorbed by the sensitive cell. Preferably, the electrical conductivity of the material making up the sensitive cell changes irreversibly when exposed to a dose of radiation.

[0032] According to another embodiment of the invention, the sensitive cell is capable of delivering an electrical signal proportional to the electrical conductivity of the material making up the sensitive cell.

[0033] According to a preferred embodiment of the invention, the sensitive cell comprises a field effect transistor. Preferably, the field effect transistor has an oxide thickness of between 80 nm and 420 nm.

[0034] Of course, the various features, variants and embodiments mentioned above may be combined with each other in various combinations, provided that they are not incompatible or mutually exclusive.

[0035] The invention also relates to a spacecraft passivation system coupled to a timing device as described above, such that the passivation system can be activated by the timing device.

[0036] Preferably, the passivation system is capable of depleting all electrical reserves on board a spacecraft.

[0037] The invention also relates to a spacecraft comprising a passivation system as described above, preferably a small satellite or nanosatellite. By nanosatellite is meant a spacecraft whose weight is between 1 kg and 50 kg, preferably between 1 kg and 30 kg. And for any satellite up to 7000 kg in GEO. Description of the figures

[0038] The invention will be better understood with regard to the detailed description below, which relates to a preferred and non-limiting embodiment, illustrated by the following figure:

[0039] [Fig. 1] illustrates a diagram of a timing device allowing the activation of a passivation system of a spacecraft, according to the invention;

[0040] [Fig. 2] illustrates a diagram of an alternative embodiment of a timing device allowing the activation of a passivation system of a spacecraft, according to the invention. Detailed description of the invention

[0041] As a reminder, the invention proposes a safer, less expensive and more compact timing device, allowing the triggering of a passivation system of a small satellite.

[0042] Figure 1 represents a schematic view of the different elements making up a timing device 2 according to the invention. Reference 4 designates a control device coupled to a dosimeter 6.

[0043] The term "coupled" means the possibility for two elements to exchange information. According to the present example, the elements making up the timing device 6 are connected to each other via electrical conductors, designated by the reference 8 in FIG. 1, allowing the exchange of information between said elements.

[0044] More specifically, the control device 4 consists of a comparison unit 10, preferably analog, coupled to a communication unit 12.

[0045] The communication unit 12 is configured to receive information from outside the timing device. The information can be transmitted to the communication unit 12 in the form of electrical signals or radiofrequency signals. Preferably, the communication unit 12 is configured to receive information from the surface of an astral body, for example a terrestrial body, when the timing device 2 is present in a spacecraft. The communication unit is also capable of processing messages internal to the satellite from an external source.

[0046] The comparison unit 10 is capable of transmitting information received by the communication unit 12 to a reference unit 14, so that said information is recorded by the reference unit. The comparison unit 10 is also capable of reading information contained in the reference unit 14, such as a reference value mentioned below. The comparison unit 10 is coupled to the dosimeter 6 so as to be able to receive, in real time or periodically, an electrical signal generated by the dosimeter 6.

[0047] According to the present example and without limitation, the dosimeter 6 comprises a field effect transistor, marketed by the company VARADIS, under one of the following names: of the following references, available on the website https: / / www.varadis.com / products / :

[0048] - Varadis RADFET VT01, comprising a 400 nm RADFET in a six-lead SOT-23 plastic package, for doses between 1 cGy (1 rad) and 1 kGy (100 krad); where

[0049] - Varadis RADFET VT03, comprising a 1 μm RADFET in a six-lead SOT-23 plastic package, for doses between 3 mGy (0.3 rad) and 10 Gy (1 krad); or

[0050] - Varadis RADFET VT05, comprising a 100 nm RADFET in a six-lead SOT-23 plastic package, for doses between 10 Gy (1 krad) and 10 kGy (1 Mrad).

[0051] Electronic components designated by the reference "Varadis RADFET" consist of an insulated gate field effect transistor, more commonly referred to as MOSFET (acronym for "Metal Oxide Semiconductor Field Effect Transistor"). More precisely, these field effect transistors have a "p" type channel, sensitive to radiation: gamma, X and protons.

[0052] The term "sensitive" means that the electrical properties of the oxide making up the field effect transistor vary depending on the dose of radiation absorbed by the oxide. In other words, the threshold voltage of the field effect transistor varies depending on the dose of radiation absorbed by the oxide forming the gate of the transistor.

[0053] The threshold voltage of the field effect transistor is biased by a direct current, for example 10 μA. Depending on the absorbed radiation dose, the value of the threshold voltage of the field effect transistor varies by an increase linearly proportional to the dose received, until reaching a saturation value defining the range of use of the sensor.

[0054] For each field effect transistor mentioned above, the semiconductor manufacturer communicates a specific calibration curve, allowing those skilled in the art to easily estimate the initial value of the threshold voltage and the electrical sensitivity (see curve) delivered by said transistor, depending on the dose of radiation it absorbs.

[0055] However, the dose of radiation absorbed by the transistor can be easily estimated using the “OMERE” software, which can be downloaded from the website of the company TRAD TESTS & RADIATIONS, at the following address: https: / / www.trad.fr / spatial / logiciel-omere / .

[0056] More precisely, the “OMERE” software makes it possible to evaluate the dose of radiation absorbed by a material, depending on its nature, its thickness as well as its time of presence in an orbit.

[0057] By selecting as a parameter for the “OMERE” software, an oxide of a nature and thickness identical to the oxide forming the “p” type channel of an insulated gate field effect transistor mentioned above, marketed by the company VARDIS, a precise orbit and a defined flight time, the person skilled in the art can easily obtain an estimate of the dose of radiation absorbed by said transistor.

[0058] By reporting the value of this estimate on the specific calibration curve of said transistor, communicated by the company VARADIS, the person skilled in the art can easily estimate the value of the electrical voltage hereinafter referred to as the reference value, delivered by said transistor at the end of a period and an orbit selected above.

[0059] The comparison unit 10 is coupled to the dosimeter 6 as well as to the reference unit 14, so as to be able to characterize in real time or periodically, the value of the electrical voltage delivered by the field effect transistor making up the dosimeter 6, then compare this electrical voltage value to a reference value pre-recorded by the reference unit 14.

[0060] Remarkably, the comparison unit 10 is configured to deliver a control signal when the value of the electrical voltage delivered by the field effect transistor making up the dosimeter 6 is equal to or exceeds the reference value recorded by the reference unit 14. The control signal is intended to allow the activation of a passivation system, not shown in the figures.

[0061] In other words, when the timing device 2 is intended to move on a known orbit, the reference value recorded by the reference unit 14 makes it possible to deterministically program the moment when the timing device 2 moving on this orbit emits a control signal capable of activating a passivation system.

[0062] Thus, advantageously, the timing device 2 evaluates the time spent in an orbit, as a function of the dose of radiation received by the dosimeter 6. Thus, the timing device 2 is more compact and safer than the electronic devices described above, because the time is characterized in relation to its environment and not as a function, for example, of a bulky electronic oscillator, the operation of which is further disturbed by the surrounding radiation.

[0063] According to a variant of the invention illustrated by figure 2, the reference unit 14 is composed of several electrical resistors 16, connected in series and powered by the same voltage source V. Preferably, the voltage source V is configured to deliver a stable current until the timing device 2 emits the passivation signal.

[0064] Advantageously, the reference unit 14 comprises two controlled switches 18 and 19 capable of being controlled by the communication unit 12. As illustrated in FIG. 2, the controlled switches 18 and 19 are mounted in parallel with one or two electrical resistors 16, so as to allow the communication unit to modify the value of the electrical voltage delivered by the reference unit 14 to the comparison unit 10.

[0065] By way of non-limiting example, when the controlled switches 18 and 19 are both held in the open position as illustrated in FIG. 2, the timing device is configured to activate a passivation device, at the end of a first flight time of the passivation device at a defined orbit.

[0066] When the communication unit 12 only controls the closing of the controlled switch 18, the timing device is reprogrammed to activate a passivation device at the end of a second flight time shorter than the first.

[0067] When the communication unit 12 commands the closing of the controlled switches 19, the timing device is then reprogrammed to immediately activate a passivation device.

[0068] Thus, the communication unit 12 advantageously makes it possible to remotely modify the desired moment when the timing device will emit a signal capable of triggering the passivation of a satellite when said device is present in a satellite in flight.

Claims

CLAIMS

1. Timing device (2) allowing the activation of a passivation system of a spacecraft, characterized in that the timing device (2) comprises a control device (4) coupled to a dosimeter (6), the dosimeter is capable of characterizing a dose of radiation to which a spacecraft is exposed during its mission, so that the control device (4) can become aware of the dose of radiation absorbed by said dosimeter (6), and in that the control device (4) is capable of activating a passivation system of a spacecraft when the dose of radiation received by said dosimeter (6) exceeds a reference value.

2. Timing device according to claim 1, characterized in that the control device (4) comprises a comparison unit (10) connected to the dosimeter (6) and a reference unit (14) connected to the comparison unit (10).

3. Timing device according to claim 2, characterized in that the reference unit (14) comprises analog components (16).

4. Timing device according to claim 3, characterized in that the reference unit (14) comprises one or more electrical resistors (16) connected in series, powered by a voltage source (V), so as to deliver to the comparison unit (10) an electrical voltage corresponding to the reference value.

5. Timing device according to claim 4, characterized in that the reference unit (14) comprises one or more controlled switches (18, 19), connected in parallel with one or more electrical resistors (16), so as to modify the electrical voltage delivered by the reference unit (14) to the comparison unit (10).

6. Timing device (2) according to one of claims 2 to 5, characterized in that the control device (4) comprises a communication unit (12) capable of receiving news. reference value and to substitute the reference value recorded by the reference unit (14) by the new reference value.

7. Timing device (2) according to claim 6, characterized in that the communication unit (12) is able to receive a new reference value when the timing device (2) is present in a spacecraft in flight.

8. Timing device (2) according to one of claims 1 to 7, characterized in that the dosimeter (6) comprises a sensitive cell, composed of a material whose physical properties change proportionally to the dose of radiation, absorbed by the sensitive cell.

9. Timing device (2) according to claim 8, characterized in that the material constituting the sensitive cell is able to absorb a dose of radiation whose value is between 1 krad and 1 Mrad.

10. Timing device (2) according to claim 8 or 9, characterized in that the electrical conductivity of the material constituting the sensitive cell changes proportionally to the dose of radiation absorbed by the sensitive cell.

11. Timing device (2) according to claim 8, characterized in that the sensitive cell is capable of delivering an electrical signal proportional to the electrical conductivity of the material constituting the sensitive cell.

12. Timing device (2) according to one of claims 8 to 11, characterized in that the sensitive cell comprises a field effect transistor.

13. Passivation system of a spacecraft characterized in that the passivation system is coupled to a timing device (2) according to one of claims 1 to 12, so that the passivation system can be activated by the timing device (2).

14. A spacecraft passivation system according to claim 13, characterized in that the passivation system is capable of depleting all electrical reserves on board a spacecraft.

15. A spacecraft comprising a passivation system according to claim 13 or 14.

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