Method and device for deorbiting an artificial satellite from Earth orbit

Abstract

The present invention relates to a method and device for deorbiting artificial satellites from Earth orbit, the artificial satellite comprising a multilayer insulation. Within the scope of the invention, it is proposed that at least one layer of the multilayer insulation be detached, at least partially, from the artificial satellite. Detaching the layer exposes the satellite structure beneath it to environmental influences. This accelerates the self-disintegration of the artificial satellite, thereby reducing its mass and increasing its ballistic coefficient. The separation of the layer increases its cross-sectional area and leads to premature reentry into the Earth's atmosphere by reducing energy. Several layers can be arranged so that, regardless of the satellite's rotation, there is always at least one surface oriented against the aerodynamic flow.

Description

Description Title of the invention: Method and device for deorbiting a artificial satellite in Earth orbit

[0001] — The present invention relates to a method and device for deorbiting a artificial satellite in Earth orbit, the artificial satellite comprising insulation multilayer.

[0002] — Multilayer insulation is used for the thermal insulation of almost all Artificial satellites. They cover the exterior of the satellites to protect them, thus than their subsystems, from the influences of the orbiting environment. They consist of made of several layers of thin, reflective, coated plastic material of metal that are separated by poorly conductive filling materials or insulators. For Earth satellites, 15 to 20 layers are generally used. (Flegel, SK, Gelhaus, J., Mockel, M., Wiedemann, C., Krag, H., Klinkrad, H., & Värsmann, P. (2011). Multi-layer insulation model for MASTER2009. Germany: Elsevier). These layers can have a total thickness of 5 to 12 mm. Such re- Clothing, for example, is known from Gabryel, W., Hoidn, W., & Wolf, J. (2016). Al- ternative Bonding Methods for MLI blankets, 46th International Conference on Environment ronmental Systems (p. 121). Austria: ICES and Finckenor, M., & Dooling, D. (1999). Multilayer Insulation Material Guidelines. Alabama: NASA Marshall Space Flight.

[0003] There are 34,000 objects larger than 10 cm, 900,000 objects smaller than- larger than 1 cm and 128 million objects larger than 1 mm in orbit Low Earth Orbit (LEO). These objects move at speeds greater than 11 km / s. 26,600 of these objects have been identified according to the ESA (ESA Space Debris Office, 2020). Each satellite launch adds at least two objects, which can then be divide into many smaller objects. One at the time of the satellite's launch, in the form of the final stage of the rocket, and the other at the end of the satellite's life, namely the A satellite without any function of its own. Since the advent of the new space age with small satellites, the number of satellite launches has increased exponentially nentielle. During the period 2014-2019, the number of satellite launches in Low Earth orbit has increased from 2-3 satellites per week to 13-14 satellites per week (Euroconsult, 2019). In 2020, 1,000 small satellites weighing less than 500 kg were launched. This corresponds to an average of 20 satellites per week. This larger number of The presence of satellites in orbit results in an increased risk of collisions. Collisions can lead to the uncontrolled introduction of debris into Earth's orbit. forming additional objects. This debris can in turn cause collisions and generate further debris through a cascade effect. This effect is also known as Kessler syndrome. All objects without function in Earth orbit are also called space debris. The threat posed by space debris continues to grow due to the launch of several multi-satellite constellations, particularly into low Earth orbit. To ensure safe access to space in the future, the UN has introduced guidelines aimed at reducing space debris for all launches. The ISO 24113 standard aims to guarantee safe access to space. According to this standard, any non-functional satellite or object in low Earth orbit must re-enter the Earth's atmosphere or be placed in a graveyard orbit within 25 years of its operational lifespan. Devices and methods for deorbiting artificial satellites are known. Active systems for deorbiting space debris are known. In these systems, space debris is collected by a satellite launched in addition to the orbiter, then burned up in the atmosphere along with the satellite. Collection can be carried out, for example, using robotic arms, nets, or harpoons. Special targets can be attached to artificial satellites before their launch to facilitate subsequent collection. The drawback of active systems is that they require additional launches. Besides active deorbiting devices, passive deorbiting devices are also known. Passive deorbiting devices are known in the form of trailing devices. In these devices, increasing the satellite's surface area allows for a greater braking effect from atmospheric remnants in near Earth orbit. This braking effect accelerates the satellite's reentry into the atmosphere. The trailing device can, for example, take the form of a trailing sail or a deployable boom. This trailing device is activated at the end of the satellite's operational life. It is known to use electromagnetic cables to deorbit satellites. This technology uses a conductive cable to generate an electromagnetic force when the cable moves relative to the Earth's magnetic field. The drawback of existing systems is that this additional weight must be placed in Earth orbit along with the satellite, increasing the cost of the rocket launch. Furthermore, there are limitations to their applicability to small satellites due to the limited forces that can be achieved. The present invention aims to optimize the targeted deorbiting of satellites. This objective is achieved with a method for deorbiting artificial satellites from Earth orbit by the fact that at least one layer of the multilayer insulation, preferably several layers, particularly favorably all the layers are detached at least partially from the artificial satellite. Artificial satellites include, for example, rockets, rocket stages, surveillance, information or communication satellites, or rocket fragments. Partial detachment of the artificial satellite's layers allows for targeted exposure, at a specific time, to the influences of the orbiting environment. This partial detachment removes at least some of the protective layers that shield the satellite from environmental influences. Now unprotected, the satellite is decomposed by oxygen radicals, ultraviolet radiation, and extreme temperature variations. For example, the material of the artificial satellite decomposes into CO, H₂O, and various metal oxides. This material degradation triggers the satellite's self-disintegration as soon as it loses its function or its connection to Earth. A preferred embodiment of the invention consists in the detachment of at least one layer occurring at the end of the life of the artificial satellite. Advantageously, the satellite remains protected from environmental influences throughout its operational life. When the satellite reaches the end of its operational life, for example due to a fault or a scheduled shutdown, self-disintegration can be triggered in a controlled manner. This can be initiated by a signal actively sent to the satellite. It can also be automatic, for example, when a certain operating time is reached or in the event of loss of contact with a control center. For this purpose, it is helpful if at least one layer of the multi-layer insulation is at least partially peeled back. By "peeling," we mean that the layer is removed in such a way that parts of the artificial satellite are freed from the protective layers of the multilayer insulation and are thus exposed to environmental influences. The layer is not separated from the satellite, but removed from a part of the satellite's structure, for example, by rolling. Another embodiment of the invention consists of separating at least one layer of the multilayer insulation in the manner of a shutter. Advantageously, the process allows the use of parts already present on the satellite. No additional membrane is needed to increase the surface area of ​​the artificial satellite. Extending the layer increases the cross-sectional area of ​​the artificial satellite. This increased cross-sectional area further slows the artificial satellite, leading to its deorbiting. The cross-sectional area can be increased by extending the layer in a flap-like fashion. Preferably, large, flat sections of the layers are separated. The braking effect then corresponds roughly to that of a taut sail. Advantageously, no additional membrane is needed to increase the cross-sectional area. This separation allows parts of the artificial satellite to be freed from the protective layers of the multilayer insulation and thus exposed to environmental influences. Advantageously, the self-disintegration of the artificial satellite is accelerated. It is preferable to orient several widely spaced layers so that, regardless of the satellite's position, there is always at least one surface oriented against the aerodynamic flow. Alternatively, individual layers and / or parts of a layer can be peeled off and / or removed. The objective is also achieved with a device for deorbiting artificial satellites from Earth orbit by providing means to detach at least partially at least one layer of the multilayer insulation of the artificial satellite. One embodiment of such a device consists in the device being disposed under a multilayer insulation of a satellite and comprising a spring element, a heating element and an adhesive agent. The device is placed beneath the multilayer insulation. The heating element can then be designed as a thin-film heater. It can be printed on thin films of polyimide or other materials suitable for space applications. Depending on the requirements, it can be printed in any shape and size. Such a thin-film heater is described, for example, in document WO 2005 / 051042 A1. In this case, the thin-film heater consists of three thin layers of minimal thickness, for example, < 2 µm. The outer layers are electrically insulating, while the middle layer is electrically resistive. The middle layer is connected to power sources. It is attached to the satellite structure beneath the multilayer insulation using an adhesive suitable for space applications. The adhesive bonds at least one layer of the artificial satellite's multilayer insulation to the satellite structure. It may be a heat-detachable adhesive. During the period in which the satellite is performing its mission, the heating element remains inactive. As soon as the satellite loses its operational capability, for example due to loss of radio contact with the ground or a collision with space debris, the heating element is automatically connected to the solar panels or the satellite's main power source. The heat generated by the heating element must be sufficient to melt the adhesive. The multilayer insulation layer, which was previously bonded to the structure of the The satellite, held in place by the adhesive, detaches from its structure. The layer is separated from the satellite structure by the spring element. The satellite structure is thus at least partially freed from the protective multilayer insulation. The satellite is now exposed to environmental influences. These environmental influences accelerate the self-disintegration of the artificial satellite. It is advantageous to initiate the described process when the satellite is outside the Earth's shadow, so that solar radiation can be used both to generate energy and to thermally enhance the detachment process. It is part of the invention that the spring element is a flat spring. The flat spring functions similarly to a child's self-adjusting wristband. After the device is activated and at least one layer of the multi-layer insulation of the satellite structure is detached, the layer is wound by the flat spring. Advantageously, the satellite structure beneath the layer is then exposed to environmental influences. The flat spring is particularly well-suited for detaching the multi-layer insulation from non-planar areas of the satellite's surface. Finally, it is in accordance with the invention that the spring element is a torsion spring. The torsion spring allows the multilayer insulation to be pulled away from the satellite's surface, like a flap. This pulling away increases the cross-sectional area of ​​the artificial satellite. The large, flat sections of the multilayer insulation are particularly well-suited for this pulling. Advantageously, an additional membrane is not required. Instead, the multilayer insulation already applied to the satellite enhances the braking effect. The artificial satellite is thus slowed by this increased braking effect until it burns up in the Earth's atmosphere. It is also possible to use a combination of flat springs and torsion springs. This can be adapted according to the geometry of the artificial satellite. An example of an embodiment of the invention is explained in more detail below with the aid of drawings. The figures show: [Fig.1] a satellite according to the invention during the usage phase, [Fig.2] a satellite according to the invention after the use phase in a first embodiment, [Fig.3] a satellite according to the invention after the use phase in a second embodiment, [Fig. 4] a deorbiting device according to the invention in the non-activated state, [Fig. 5] a deployment mechanism according to the invention in the deactivated state, [Fig. 6] a satellite according to the invention after the use phase with a first mechanism activated [Fig.7] a satellite according to the invention after the use phase with a second mechanism activated. Figure 1 shows a satellite (1) according to the invention during its operational phase. The deorbiting device (2) is concealed beneath the multilayer insulation (3). During its operational phase, the multilayer insulation (3) protects the satellite's structure and payload from environmental influences, including radiation and temperature changes. After the satellite (1) has ceased operating, the deorbiting device (2) according to the invention modifies the multilayer insulation (3) to increase the satellite's (1) surface area. This can be achieved by peeling back the multilayer insulation (3) using a peeling mechanism or by deploying the multilayer insulation (3) using a deployment mechanism. Figure 2 shows a satellite (1) according to the invention after the use phase with the deorbiting device activated, which is here a peeling mechanism (4). When the peeling mechanism (4) is activated, the multilayer insulation (3) is rolled up and remains on the satellite (1) which is exposed to space radiation, which initiates the decomposition of the satellite structure, a reduction in mass and finally deorbiting. Figure 3 shows a satellite (1) according to the invention after the use phase with an activated deorbiting device comprising a deployment mechanism (5). After the multilayer insulation (3) is peeled away, it is extended like a flap by means of the deployment mechanism (5), thus significantly increasing the cross-sectional area of ​​the satellite (1), which greatly slows it down. Figure 4 shows in detail a peeling mechanism (4) according to the invention in the non-activated state, consisting of a heating element (6) designed as a thin film heating element, a flat spring (7) integrated into the multilayer insulation (3) and an adhesive agent (8). Figure 5 shows in detail a deployment mechanism (5) according to the invention in the non-activated state, consisting of a heating element (6) designed as a thin film heating element, an adhesive agent (8) and a torsion spring (9) pre-stressed in the multilayer insulation (3). Figure 6 shows a satellite (1) according to the invention after the use phase with a mechanism activated after the use phase, composed of the thin film heating element (6), the adhesive agent (8) and the torsion spring (9), and in which the multilayer insulation (3) is extended in the manner of a flap by the tension force of the torsion spring (9). Figure 7 shows a satellite (1) according to the invention after the use phase with a peeling mechanism activated after the use phase and composed of the flat spring (7), and in which the multilayer insulation (3) is wound by the tensile force = resso dish (7).

Claims

Demands

1. A method for deorbiting artificial satellites (1) from Earth orbit, the artificial satellite (1) comprising a multilayer insulation (3), ca- characterized in that at least one layer of the multilayer insulation (3) is detached at least partially from the artificial satellite (1).

2. A method according to claim 1, characterized in that the detachment of at least one layer occurs at the end of the life of the artificial satellite (1).

3. Method according to claim 1 or 2, characterized in that at least a layer of the multilayer insulation (3) is at least partially Peleus.

4. A method according to any one of the preceding claims, characterized in that that at least one layer of the multilayer insulation (3) is separated at the in the manner of a shutter.

5. A method according to any one of the preceding claims, characterized in that that several separated layers are oriented in such a way that, what Whatever the position of the satellite (1), there is always at least one surface oriented against the aerodynamic flow.

6. Device for deorbiting artificial satellites from Earth orbit according to a method conforming to one of the preceding claims, ca- characterized in that means (6, 7, 8, 9) are provided for detaching to less partially at least one layer of the multilayer insulation (3) of the artificial satellite.

7. Device according to claim 6, characterized in that the device is disposed under the multilayer insulation (3) of a satellite and comprises a spring element (7, 9), a heating element (6), and an adhesive agent (8).

8. Device according to claim 7, characterized in that the element spring is a flat spring (7).

9. Device according to claim 7, characterized in that the element spring is a torsion spring (9).

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