Dislocation tolerant spacecraft shielding

Abstract

The application provides a space vehicle protection device capable of dislocation, which comprises a fixed layer, a connecting rod and a rotating layer, the fixed layer is installed on the surface of a space vehicle; one end of the connecting rod is arranged on the fixed layer; the rotating layer is rotatably arranged on the other end of the connecting rod; wherein when the rotating layer and the fixed layer generate aligned impact holes, an astronaut drives the rotating layer to rotate, so that the impact holes on the rotating layer are dislocated from the impact holes on the fixed layer. In the application, when the rotating layer and the fixed layer generate aligned impact holes, the impact holes on the rotating layer are dislocated from the impact holes on the fixed layer by rotating the rotating layer, so that at least one layer of protection is ensured in front of the space vehicle, the space vehicle is prevented from being directly exposed to the space debris environment, and the protection capability of the protection device is improved without increasing additional protection structures.

Description

Technical Field

[0001] This application relates to the technical field of space debris protection, and more specifically, to a dislocationable spacecraft protection device. Background Technology

[0002] Space debris consists of waste left in space by human space activities. It mainly includes rocket final stages and nose cones, defunct satellites and spacecraft, fuel ejected from solid rockets, and debris from spacecraft collisions. With the rapid development of space technology, the number of spacecraft in orbit continues to surge, making the space debris environment increasingly harsh. Space debris ranging in size from micrometers to centimeters possesses extremely high orbital speeds and enormous impact kinetic energy, posing a core threat to the safety of spacecraft in orbit.

[0003] When space debris collides with a spacecraft, the average collision speed can reach 10 km / s. Even space debris the size of millimeters can cause serious damage to critical components of a spacecraft. After being hit by space debris, a spacecraft may experience changes in its bulkhead performance, surface pitting, perforation, or even serious accidents such as container rupture and explosion, posing a serious threat to the spacecraft's safety in space.

[0004] In existing technologies, spacecraft space debris protection structures are mostly fixed passive protection structures, including single-layer protective plates, double-layer whipple structures, multi-layer sandwich protective structures, and honeycomb protective structures. Their core protection principle is to block and break high-speed debris through protective plates, thereby consuming the impact kinetic energy of the debris and achieving spacecraft protection.

[0005] However, existing protective structures are fixed and, after being subjected to high-speed impacts from space debris, permanent, non-healing through-holes form on the surface. This fails to fundamentally solve the problem of secondary protection failure after penetration, leaving the protective structure with the inherent defect of permanent failure after a single instance of damage, resulting in low protection lifespan and reliability. During the long-term operation of a spacecraft in orbit, if subsequent space debris enters along the path of this through-hole, the area has no protective capabilities and can directly penetrate the protective structure to impact the spacecraft body, causing damage to the spacecraft shell, equipment failure, and even triggering major safety accidents such as on-orbit disintegration and leakage. Summary of the Invention

[0006] The purpose of this application is to provide a dislocation-resistant spacecraft protection device that can solve at least one of the aforementioned technical problems. The specific solution is as follows: According to a specific embodiment of this application, this application provides a dislocationable spacecraft protection device, which is installed on the surface of a spacecraft. The spacecraft protection device includes a fixed layer, a connecting rod, and a rotating layer. The fixed layer is installed on the surface of the spacecraft. One end of the connecting rod is disposed on the fixed layer. The rotating layer is rotatably disposed on the other end of the connecting rod. When the rotating layer and the fixed layer generate aligned impact holes, the astronaut drives the rotating layer to rotate, causing the impact holes on the rotating layer to be dislocated from the impact holes on the fixed layer.

[0007] In some embodiments, the spacecraft protection device further includes a rotating assembly, the rotating assembly including a rotating shaft, the main body of the rotating shaft passing through the connecting rod, and one end of the rotating shaft being connected to the rotating layer.

[0008] In some embodiments, the rotating assembly further includes a rotating cylinder rotatably disposed on the other end of the connecting rod, the rotating layer being rotatably disposed on the rotating cylinder, and the rotating cylinder being connected to the rotating shaft.

[0009] In some embodiments, the rotating layer and the fixed layer have the same surface dimensions, and the surfaces of the rotating layer and the fixed layer are provided with the same marking layer.

[0010] In some embodiments, the surface of the marking layer is provided with dividing lines; preferably, the number of dividing lines is 6 or 12, and the angle between adjacent dividing lines is 60° or 30°.

[0011] In some embodiments, the material of the rotating layer includes, but is not limited to, a metal plate, an alloy plate, or a composite material plate.

[0012] In some embodiments, the material of the fixing layer includes, but is not limited to, a fiber cloth woven material board, a foam board, or an aluminum alloy board.

[0013] In some embodiments, the rotating layer and the fixing layer are square plates or regular hexagons.

[0014] In some embodiments, the spacing between the rotating layer and the fixed layer is less than or equal to the diameter of the rotating layer.

[0015] Compared with the prior art, the above-described solutions of this application have at least the following beneficial effects: In the dislocationable spacecraft protection device of this application, after the rotating layer and the fixed layer produce aligned impact holes, the rotating layer is rotated to displace the impact holes on the rotating layer from the impact holes on the fixed layer, so that the protection device can continue to be used. This ensures that there is at least one layer of protection in front of the spacecraft, preventing the spacecraft from being directly exposed to the space debris environment, and improving the protection capability of the protection device without adding additional protection structures. Attached Figure Description

[0016] The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments of the present application and, together with the description, serve to explain the principles of the application. It is obvious that the drawings described below are merely some embodiments of the present application, and those skilled in the art can obtain other drawings based on these drawings without any inventive effort. In the drawings: Figure 1 A schematic diagram of a dislocation-enabled spacecraft protection device provided in an embodiment of the present invention.

[0017] Figure 2 This is a schematic diagram of a marker layer in a dislocationable spacecraft protection device provided in an embodiment of the present invention.

[0018] Figure 3 Another schematic diagram of the marker layer in a dislocationable spacecraft protection device provided in an embodiment of the present invention.

[0019] Explanation of reference numerals in the attached figures: Spacecraft protective device 100, fixed layer 10, connecting rod 20, rotating layer 30, rotating component 40, rotating shaft 41, rotating cylinder 42, marking layer 50, and dividing line 51. Detailed Implementation

[0020] To make the objectives, technical solutions, and advantages of this application clearer, the application will be further described in detail below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, and not all embodiments. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0021] It should also be noted that the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that an article or device that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such an article or device. Without further limitation, an element defined by the phrase "comprising one" does not exclude the presence of other identical elements in the article or device that includes said element.

[0022] According to relevant technologies, there are currently more than 54,000 pieces of space debris larger than 10 cm, more than 1.2 million pieces of space debris between 1 and 10 cm, more than 140 million pieces of space debris between 1 mm and 1 cm, and more than 10 billion pieces of space debris smaller than 1 mm. When low-Earth orbit space debris collides with spacecraft, the average collision velocity can reach 10 km / s. Even millimeter-sized space debris can cause serious damage to critical components of spacecraft. After being hit by space debris, spacecraft may experience changes in bulkhead performance, surface craters, perforations, or even container rupture and explosions, posing a serious threat to the safety of spacecraft in space.

[0023] The prior art CN103466104B discloses a composite SiC fabric-filled space debris protection structure. The composite SiC fabric prepared by processes such as weaving, impregnation, and heat treatment replaces the foreign Nextel fiber fabric in the preparation of a filled protective structure for the protection of spacecraft space debris.

[0024] The prior art patent CN107140238A discloses a kinetic energy efficient dissipation space debris protection structure, which is processed into a kinetic energy efficient dissipation protection screen by welding or adhesively connecting high-performance wave impedance gradient materials, porous materials and back plates, giving full play to the advantages of each layer of materials and realizing the lightweight of the protection structure.

[0025] The prior art CN 110155375 B discloses a Whipple-type high-performance space debris protection structure, which combines the inner support layer and the outer protection layer together by adopting a honeycomb support structure to form a double-layer barrel-shaped protection structure, thereby achieving stress dispersion and protection for large hulls.

[0026] However, existing technologies have the following shortcomings: the protective structure is a fixed structure, and after being subjected to high-speed impacts from space debris, permanent, non-healing through holes will form on the surface; traditional structures cannot fundamentally solve the problem of secondary protection failure after perforation, and the protective structure still suffers from the defect of permanent failure after a single damage, resulting in low protection lifespan and reliability. During the long-term on-orbit service of a spacecraft, if subsequent space debris enters along the path of the through hole, the location of the through hole has no protective capability, and the debris can directly penetrate the protective structure and impact the spacecraft body, causing damage to the spacecraft shell, equipment failure, and even triggering major safety accidents such as on-orbit disintegration and leakage.

[0027] This application addresses the technical pain point of existing fixed space debris protection structures failing to provide secondary impact protection after impact perforation. It proposes a spacecraft protection device that allows for the displacement of space debris. This device can be repaired without the need for additional protection structures and can be moved or rotated, effectively solving the defect of permanent failure after a single damage and improving the reliability of secondary impact protection.

[0028] The following is in conjunction with the appendix Figures 1-3 Detailed description of optional embodiments of the present invention.

[0029] like Figure 1 As shown, according to a specific embodiment of this application, this application provides a dislocationable spacecraft protection device 100, which is installed on the surface of a spacecraft. The dislocationable spacecraft protection device 100 includes a fixed layer 10, a connecting rod 20, and a rotating layer 30.

[0030] The fixed layer 10 is installed on the surface of the spacecraft; one end of the connecting rod 20 is disposed on the fixed layer 10; the rotating layer 30 is rotatably disposed on the other end of the connecting rod 20; wherein, when the rotating layer 30 and the fixed layer 10 produce aligned impact holes, the astronaut drives the rotating layer 30 to rotate, causing the impact holes on the rotating layer 30 to be misaligned with the impact holes on the fixed layer 10.

[0031] In this embodiment, the working mechanism of the rotating layer 30 is that the collision between the space debris and the rotating layer 30 generates an ultra-high pressure shock wave of tens of GPa. Simultaneously, the space debris melts, vaporizes, and pulverizes at the contact area with the rotating layer 30, and the entire debris is broken into a micron-sized solid-liquid-gas mixed debris cloud. The connecting rod 20 is used to connect the two protective layers, allowing the fragmented debris cloud to freely diffuse in all directions within the distance between the fixed layer 10 and the rotating layer 30. This spreads the impact momentum, which was originally concentrated at a single point, over a larger area, while also lengthening the impact time and reducing the instantaneous peak pressure of the fixed layer 10.

[0032] In this embodiment, the fixing layer 10 is installed on the external structural surface of the spacecraft using high-strength bolts or a special adhesive. Its surface is precision-machined to ensure a flatness error of no more than ±0.02 mm. One end of the connecting rod 20 is fixed to a pre-set mounting hole in the fixing layer 10 via an interference fit or welding. The connecting rod 20 is made of titanium alloy TC4, which has high rigidity and fatigue resistance. The rotating layer 30 is rotatably mounted on the other end of the connecting rod 20. Initially, neither the rotating layer 30 nor the fixing layer 10 has impact holes. When the rotating layer 30 and the fixing layer 10 collide with space debris, if aligned impact holes are formed, the rotating layer 30 is rotated to misalign the impact holes on the rotating layer 30 with those on the fixing layer 10. This ensures at least one layer of protection in front of the spacecraft, preventing direct exposure to the space debris environment and enhancing the protective capability of the dislocation-prone spacecraft protection device without adding additional protective structures.

[0033] In some embodiments, such as Figure 1As shown, the dislocationable spacecraft protection device 100 further includes a rotating assembly 40, which includes a rotating shaft 41. The rotating shaft 41 passes through the connecting rod 20, and one end of the rotating shaft 41 is connected to the rotating layer 30. In this embodiment, the rotating assembly 40 serves as the core power and transmission unit of the dislocationable spacecraft protection device, and its design must meet the reliability and stability requirements of the spacecraft in the extreme environment of space. The rotating shaft 41 is mostly made of high-strength alloy or carbon fiber composite material, possessing excellent fatigue strength and lightweight characteristics. Its structural design, which passes through the connecting rod 20, not only achieves smooth transmission of rotational motion but also provides necessary radial and axial support for the rotating shaft 41 through the connecting rod 20, preventing eccentricity or resonance during high-speed rotation. The rotating layer 30 is rigidly connected to one end of the rotating shaft 41 (e.g., by interference fit or high-strength bolt fixation) to ensure synchronous movement during rotation.

[0034] In some embodiments, such as Figure 1 As shown, the rotating assembly 40 also includes a rotating cylinder 42, which is rotatably mounted on the other end of the connecting rod 20. The rotating layer 30 is rotatably mounted on the rotating cylinder 42, and the rotating cylinder 42 is connected to the rotating shaft 41. In this embodiment, the rotating cylinder 42 is typically a hollow cylindrical structure with bearings or sliding fit structures on its inner wall to ensure stable rotational connection with the other end of the connecting rod 20 and reduce frictional resistance and energy loss during rotation. The rotating shaft 41, as the core transmission component of the rotating assembly 40, transmits power to the rotating cylinder 42 through key connection, threaded connection, or interference fit. The rotation of the rotating shaft 41 drives the rotating cylinder 42 and the rotating layer 30 to rotate synchronously. This structural design not only realizes the independent rotation function of the rotating layer 30 but also ensures the stability and positioning accuracy of the entire rotating assembly 40 during operation through the connection of the connecting rod 20 to the main structure. In addition, the material selection of the rotating cylinder 42 usually considers its strength, wear resistance, and corrosion resistance. Common materials include stainless steel, high-strength alloys, or engineering plastics to adapt to different working environments and load requirements.

[0035] In some embodiments, such as Figures 2-3 As shown, the rotating layer 30 and the fixed layer 10 have the same surface dimensions, and both surfaces are provided with identical marking layers 50. In this embodiment, the rotating layer 30 and the fixed layer 10 have the same surface dimensions to ensure precise alignment during assembly and use, reducing errors caused by dimensional differences. The marking layers 50 on the surfaces of the rotating layer 30 and the fixed layer 10 are typically made of high-contrast materials, such as coatings of different colors or reflective materials, to facilitate visual identification and positioning. The design of the marking layers 50 matches the functional requirements of the rotating layer 30 and the fixed layer 10, and the marking layers 50 serve as a reference for alignment.

[0036] In some embodiments, such as Figures 2-3 As shown, the surface of the marking layer 50 is provided with dividing lines 51; preferably, the number of dividing lines 51 is 6 or 12, and the angle between adjacent dividing lines 51 is 60° or 30°. In this embodiment, the function of the dividing lines 51 is to divide the surface of the marking layer 50 into multiple uniform fan-shaped areas or equally divided areas, facilitating alignment operations. When the number of dividing lines 51 is 6, the angle between adjacent dividing lines is 60°. This design is often used in scenarios where a circle needs to be divided into 6 equal parts. When the number of dividing lines 51 is 12, the angle between adjacent dividing lines is 30°. The shape of the dividing lines 51 can be a straight line, an arc, or a specific pattern, and its depth and width need to be optimized according to the requirements of the actual application.

[0037] In some embodiments, the material of the rotating layer 30 includes, but is not limited to, metal plates, alloy plates, or composite material plates. The working mechanism of the rotating layer 30 is that the collision between the space debris and the rotating layer 30 generates an ultra-high pressure shock wave of tens of GPa. The space debris and the rotating layer 30 undergo impact melting, vaporization, and pulverization, and the entire debris is broken down into a micron-sized solid-liquid-gas mixed debris cloud.

[0038] In some embodiments, the material of the fixing layer 10 includes, but is not limited to, a woven fiber material board, a foam board, or an aluminum alloy plate. The fixing layer 10 absorbs the impact kinetic energy of the debris cloud to prevent the spacecraft behind the protective layer from being impacted or to reduce the impact energy.

[0039] In some embodiments, the rotating layer 30 and the fixed layer 10 are square plates or regular hexagons. In a specific embodiment of this application, the rotating layer 30 is made of aviation aluminum alloy plate, with a size of 30cm × 30cm and a thickness of 1.5mm, and a central opening in the square protective plate. The connecting rod 20 is made of aviation aluminum alloy plate, with a length of 10cm and a diameter of 1cm, and is a rod-shaped structure. One end of the connecting rod 20 is connected to the rotating layer 30, allowing the rotating layer 30 to rotate around the structural axis of the connecting rod 20. The fixed layer 10 is made of foam aluminum plate, with a size of 30cm × 30cm and a thickness of 10mm, and a central opening in the square protective plate. The other end of the connecting rod 20 is connected and fixed to the center of the fixed layer 10. When the dislocationable spacecraft protection device 100 in this application is subjected to a vertical impact from large-sized space debris, a Cartesian coordinate system is established with the center of the rotating layer 30. Assuming the impact point coordinates are (0, 10cm), the center coordinates of the impact holes in both the rotating layer 30 and the fixed layer 10 are (0, 10cm) and (0, 10cm), respectively. The spacecraft is directly exposed to the debris environment through the through-hole, increasing the impact risk. By rotating the rotating layer 30 by 90° or 180°, the impact hole coordinates of the rotating layer 30 become (10cm, 0) or (0, -10cm), and the impact hole coordinates of the fixed layer 10 become (0, 10cm). The perforation positions of the front and rear protective layers are staggered, ensuring that there is at least one protective structure in front of the spacecraft structure, preventing the spacecraft from being directly exposed to the space debris environment, and improving the protective capability of the protective structure without adding an additional protective layer.

[0040] In some embodiments, the spacing between the rotating layer 30 and the fixed layer 10 is less than or equal to the diameter of the rotating layer 30.

[0041] Compared with the prior art, the above-described solutions of this application have at least the following beneficial effects: In the dislocationable spacecraft protection device of this application, after the rotating layer and the fixed layer produce aligned impact holes, the rotating layer is rotated to displace the impact holes on the rotating layer from the impact holes on the fixed layer. This allows the dislocationable spacecraft protection device to continue to be used, ensuring that there is at least one layer of protection in front of the spacecraft and preventing the spacecraft from being directly exposed to the space debris environment. This enhances the protection capability of the dislocationable spacecraft protection device without adding additional protective structures.

[0042] The core logic of traditional protective structure failure is: impact penetration → permanent hole presence → secondary debris penetration along the hole → protection failure. This application breaks away from the traditional passive protection mindset. It eliminates the need for repairing or sealing the penetration hole, and requires no new protective structure. Instead, it alters the effective protection area of ​​the protective plate through structural misalignment, removing the failed, perforated area from the protective function and allowing intact plates to take over. This approach completely solves the problem of secondary protection failure with minimal structural cost and zero load increase, achieving the reuse and long-term protection of the protective structure.

[0043] This application fundamentally solves the problem of secondary protection failure: addressing the core pain point of permanent failure after the traditional protective structure is punctured, it uses a misalignment adjustment method to shift the through hole of the protective structure formed by the impact, completely avoiding the risk of secondary fragments penetrating along the hole, thus solving the problem of protection failure caused by single impact damage in principle and greatly improving the reliability of protection.

[0044] This application features zero additional load and low-cost upgrades: It does not require thickening of the protective plate, adding of multiple layers of protective structure, or replacement of the main protective material. It achieves functional upgrades entirely based on the original protective plate, without increasing the spacecraft's on-orbit load or requiring significant modifications to the existing protective system. It is highly adaptable and has low application costs.

[0045] This application provides long-lasting cyclic protection with a long service life: it can achieve repeated misalignment adjustment after multiple impacts, and the protective capability can be restored through attitude fine-tuning whether it is a single puncture or multiple punctures. It completely changes the drawback of traditional protective structures that are "one-time damage, permanent failure" and greatly extends the service life of spacecraft protection systems.

[0046] Finally, it should be noted that the various embodiments in this specification are described in a progressive manner, with each embodiment focusing on its differences from other embodiments. Similar or identical parts between embodiments can be referred to interchangeably. For the systems or apparatus disclosed in the embodiments, since they correspond to the methods disclosed in the embodiments, the descriptions are relatively simple, and relevant parts can be referred to the method section.

[0047] The above embodiments are only used to illustrate the technical solutions of this application, and are not intended to limit them. Although this application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features. Such modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of this application.

Claims

1. A dislocation-compatible spacecraft protection device, installed on the surface of a spacecraft, characterized in that, The spacecraft protection device includes: A fixing layer, which is mounted on the surface of the spacecraft; A connecting rod, one end of which is disposed on the fixed layer; A rotating layer, which is rotatably disposed on the other end of the connecting rod; When the rotating layer and the fixed layer produce aligned impact holes, the astronaut drives the rotating layer to rotate, causing the impact holes on the rotating layer to be misaligned with the impact holes on the fixed layer.

2. The spacecraft protection device capable of dislocation according to claim 1, characterized in that, Also includes: A rotating assembly, comprising a rotating shaft, the main body of which passes through the connecting rod, and one end of which is connected to the rotating layer.

3. The spacecraft protection device capable of dislocation according to claim 2, characterized in that, The rotating assembly further includes a rotating cylinder, which is rotatably disposed on the other end of the connecting rod. The rotating layer is rotatably disposed on the rotating cylinder, and the rotating cylinder is connected to the rotating shaft.

4. The spacecraft protection device capable of dislocation according to claim 1, characterized in that, The rotating layer and the fixed layer have the same surface dimensions, and the surfaces of the rotating layer and the fixed layer are provided with the same marking layer.

5. The spacecraft protection device capable of dislocation according to claim 4, characterized in that, The surface of the marking layer is provided with dividing lines; Preferably, the number of the dividing lines is 6 or 12, and the angle between adjacent dividing lines is 60° or 30°.

6. The spacecraft protection device capable of dislocation according to claim 1, characterized in that, The material of the rotating layer includes, but is not limited to, metal plates, alloy plates, or composite material plates.

7. The spacecraft protection device capable of dislocation according to claim 1, characterized in that, The material of the fixing layer includes, but is not limited to, woven fiber cloth board, foam board or aluminum alloy board.

8. The spacecraft protection device capable of dislocation according to claim 4, characterized in that, The rotating layer and the fixed layer are square plates or regular hexagonal plates.

9. The spacecraft protection device capable of dislocation according to claim 8, characterized in that, The distance between the rotating layer and the fixed layer is less than or equal to the diameter of the rotating layer.

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