A shock-resistant aircraft structure with secondary vibration reduction

Abstract

This invention relates to a high-impact resistant aircraft structure with two-stage vibration damping, comprising: an aircraft structure body housing multiple internal components; multiple first-stage vibration dampers fitted onto the aircraft structure body and configured to reduce the impact overload force on the aircraft structure body through their own compressive deformation; multiple second-stage vibration dampers disposed between the multiple internal components and the aircraft structure body, further reducing the impact overload force on the internal components; and a drive component in close contact with the first-stage vibration dampers and configured to bear the impact overload force and transmit the impact overload force to the first-stage vibration dampers. The high-impact resistant aircraft structure of this invention incorporates first-stage and second-stage vibration dampers, achieving a high attenuation ratio vibration reduction for near-ten-thousand-g level impacts in the time domain through two-stage composite vibration damping, protecting the normal operation of the individual aircraft components.

Description

Technical Field

[0001] This invention relates to the field of aircraft technology, and in particular to an aircraft structure with secondary vibration reduction to resist large impacts. Background Technology

[0002] Currently, the common solution for space-based systems to address the issue of sensitive individual units being unable to withstand vibration excitation and static overload during the active phase of the launch vehicle is to add vibration damping components at the unit's installation point to absorb vibration and reduce energy, thereby protecting the unit. However, conventional unit-level vibration damping measures cannot cope with instantaneous large overload impacts on the order of nearly 10,000 g in the time domain, and conventional space-based vibration damping methods cannot achieve a large attenuation ratio when facing high-magnitude impacts.

[0003] Vibration reduction methods with high attenuation ratios for small space-based aircraft are currently scarce. Ground-based shock absorbers designed to withstand large overload impacts have too large a spatial envelope and mass, making them unsuitable for small space-based systems. Furthermore, ground-based shock absorbers with high overload impacts have a uniform configuration, resulting in poor compatibility with aircraft structures and making them unsuitable for use on aircraft. Summary of the Invention

[0004] To address at least some of the aforementioned problems in the prior art, the present invention aims to provide a high-impact aircraft structure with secondary vibration reduction, comprising:

[0005] The spacecraft's main structure contains multiple internal satellite units;

[0006] Multiple primary shock absorbers are fitted onto the aircraft structure and configured to reduce the impact overload force on the aircraft structure through their own compressive deformation.

[0007] Multiple secondary vibration dampers are disposed between the multiple intra-satellite units and the spacecraft structure to further reduce the impact overload force on the intra-satellite units; and

[0008] The drive component is in close contact with the first-stage damper and is configured to withstand and transmit the impact overload force to the first-stage damper.

[0009] Furthermore, the aircraft structure body is a one-piece molded structure, and the aircraft structure body includes:

[0010] The main body of the spacecraft is hollow inside, and multiple onboard units are installed inside the main body of the spacecraft.

[0011] A front section structure, located at the front end of the main body of the aircraft, and the front section structure is hollow inside;

[0012] A first vibration damping mounting structure, located between the front structure and the aircraft body, is configured to provide a mounting position for the first-stage vibration damper and prevent the first-stage vibration damper from sliding under impact overload forces; and

[0013] The second vibration damping mounting structure is located at the end of the aircraft body and is configured to provide a mounting position for the first-stage vibration damper and prevent the first-stage vibration damper from sliding when subjected to impact overload force.

[0014] Furthermore, the first vibration damping mounting structure includes:

[0015] The first main body is configured to provide a mounting position for the first stage vibration damper;

[0016] A first blocking portion, located at one end of the first main body portion, protrudes outward relative to the first main body portion and is configured to prevent the first-stage shock absorber from sliding under impact overload force; and

[0017] A first mounting portion is located at the other end of the first main body portion and is connected to the drive component; and / or

[0018] The second vibration damping mounting structure includes:

[0019] The second main body is configured to provide a mounting position for the first-stage vibration damper; and

[0020] The second blocking part, located at one end of the second main body, protrudes outward relative to the second main body and is configured to prevent the first-stage shock absorber from sliding when subjected to impact overload force.

[0021] Furthermore, the plurality of first-stage vibration damper sleeves are annular vibration dampers, which are integrally molded and fitted onto the first vibration damping mounting structure and the second vibration damping mounting structure, including:

[0022] A metal plate of uniform wall thickness is heated and pressed into a corrugated shape using a mold, then wound into a ring, and finally welded together to form a single unit, thus obtaining the first-stage vibration damper.

[0023] Furthermore, the first-stage damper comprises N / 2 corrugated units, where N≥3; and / or

[0024] All edges of the first-stage vibration damper are rounded.

[0025] Furthermore, the driving component includes a first driving component and a second driving component, wherein the first driving component surrounds the first vibration damping mounting structure and a first-stage vibration damper mounted on the first vibration damping mounting structure, and the second driving component surrounds the second vibration damping mounting structure and a first-stage vibration damper mounted on the second vibration damping mounting structure. Both the first driving component and the second driving component have a bearing boss on their inner sides, and the bearing boss is in close contact with the first-stage vibration damper.

[0026] Furthermore, it also includes:

[0027] A positioning boss is configured to provide radial positioning for the drive component and to cooperate with the first-stage damper to fix the drive component.

[0028] Furthermore, the positioning boss includes:

[0029] A first positioning boss is sleeved on the first vibration damping mounting structure and fixedly connected to the first mounting part. The first positioning boss includes a first radial positioning part and a first bearing part, wherein the outer diameter of the first bearing part is larger than the outer diameter of the first radial positioning part; and / or

[0030] The second positioning boss is installed at the end of the second vibration damping mounting structure. The second positioning boss includes a second radial positioning part and a second bearing part, wherein the outer diameter of the second bearing part is larger than the outer diameter of the second radial positioning part.

[0031] Furthermore, the first driving component is sleeved on the first positioning boss, the bearing boss of the first driving component is located on the first bearing part of the first positioning boss and below the first stage vibration damper, and the inner surface of the bearing boss of the first driving component is in contact with the outer surface of the first radial positioning part of the first positioning boss.

[0032] Furthermore, it also includes a gasket, which is sleeved on the second positioning boss and located between the second bearing portion and the bearing boss of the second driving component.

[0033] Furthermore, the second driving component is sleeved on the second positioning boss, the bearing boss of the second driving component is located above the pad and below the first-stage damper, and the inner surface of the bearing boss of the second driving component is in contact with the outer surface of the second radial positioning part of the second positioning boss.

[0034] Furthermore, the multiple intra-satellite units are fastened to the inner wall of the spacecraft body by bolts, and each bolt is fitted with a second-stage vibration damper, wherein the second-stage vibration damper is a silicon-based rubber vibration damper or a metal-rubber vibration damper.

[0035] The present invention has at least the following beneficial effects:

[0036] (1) The impact-resistant aircraft structure of the present invention is provided with a first-stage vibration damper and a second-stage vibration damper. The first-stage vibration damper is sleeved on the aircraft structure body and reduces the impact overload force on the aircraft structure body through its own compression deformation. The second-stage vibration damper is set between the satellite unit and the aircraft structure body, and can further reduce the impact overload force on the satellite unit. Through the two-stage composite vibration damping, the time domain can achieve a level of nearly 10,000 g (100,000 m / s). 2 High attenuation ratio for shock absorption (level 1) to protect the normal operation of individual aircraft.

[0037] (2) The first-stage shock absorber of the present invention participates in the transmission chain of the drive component and the spacecraft structure to resist instantaneous large impacts, which can effectively reduce the force level of the spacecraft structure and the internal satellite unit installed inside, and provide protection for the spacecraft structure and the internal satellite unit.

[0038] (3) The first-stage vibration damper is a structural-level ring vibration damper. Multiple sets of first-stage vibration dampers are placed on the outside of the front end of the aircraft structure, which can effectively improve the energy storage capacity and structural stability of the vibration damper, while providing the aircraft with higher lateral stiffness.

[0039] (4) The first-stage shock absorber is a ring and is fitted onto the aircraft structure body by an integral molding method, which has good compatibility with the aircraft structure body.

[0040] (5) The overall mass of the first-stage and second-stage shock absorbers of the present invention is lower than that of ground shock absorbers with the same performance that resist large overload impacts, and they occupy less space.

[0041] (6) The aircraft structure of the present invention adopts a two-stage vibration reduction scheme, which sets up a first-stage vibration damper and a second-stage vibration damper, and can achieve a large attenuation ratio vibration reduction for time-domain impacts of nearly 10,000 g under the premise of limited space and limited mass.

[0042] (7) Compared with conventional single-machine level vibration reduction, the aircraft structure of the present invention adopts two-stage vibration reduction, which can cope with a large number of impact overload forces. Attached Figure Description

[0043] To further illustrate the above and other advantages and features of the various embodiments of the present invention, a more specific description of the embodiments of the invention will be presented with reference to the accompanying drawings. It is to be understood that these drawings depict only typical embodiments of the invention and are therefore not intended to limit its scope. In the drawings, identical or corresponding parts will be indicated by identical or similar reference numerals for clarity.

[0044] Figure 1A cross-sectional structural schematic diagram of a high-impact resistant aircraft structure according to an embodiment of the present invention is shown.

[0045] Figure 2 A schematic diagram of the structure of a first-stage vibration damper according to an embodiment of the present invention is shown.

[0046] Figure 3 A cross-sectional structural schematic diagram of a first-stage vibration damper according to an embodiment of the present invention is shown.

[0047] Figure 4 A schematic diagram of the two-stage shock absorption principle of a high-impact aircraft structure according to the present invention is shown. Detailed Implementation

[0048] It should be noted that the components in the accompanying drawings may be shown exaggerated for illustrative purposes and may not be to scale.

[0049] In this invention, the various embodiments are merely intended to illustrate the solutions of the invention and should not be construed as limiting.

[0050] In this invention, unless otherwise specified, the quantifiers “a” and “one” do not exclude scenarios involving multiple elements.

[0051] It should also be noted that, in the embodiments of the present invention, only a portion of the parts or components may be shown for clarity and simplicity. However, those skilled in the art will understand that, under the teachings of the present invention, the required parts or components can be added as needed for specific scenarios.

[0052] It should also be noted that within the scope of this invention, the terms "same", "equal", and "equal to" do not mean that the two values ​​are absolutely equal, but allow for a certain reasonable error. In other words, the terms also cover "substantially the same", "substantially equal", and "substantially equal to".

[0053] It should also be noted that in the description of this invention, the terms "center," "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for simplifying the description, and do not explicitly or implicitly suggest that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on the invention. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0054] Furthermore, the embodiments of the present invention describe the process steps in a specific order. However, this is only for the convenience of distinguishing each step, and is not a limitation on the order of each step. In different embodiments of the present invention, the order of each step can be adjusted according to the process.

[0055] In this invention, the term "configuration" refers to setting the shape, structure, material and / or function of a target object to achieve a desired technical effect. "Configuration" includes a variety of alternative technical means to achieve the technical effect, which become apparent under the teachings of this invention.

[0056] Figure 1 A cross-sectional structural schematic diagram of a high-impact resistant aircraft structure according to an embodiment of the present invention is shown. Figure 2 A schematic diagram of the structure of a first-stage vibration damper according to an embodiment of the present invention is shown. Figure 3 A cross-sectional structural schematic diagram of a first-stage vibration damper according to an embodiment of the present invention is shown. Figure 4 A schematic diagram of the two-stage shock absorption principle of a high-impact aircraft structure according to the present invention is shown.

[0057] like Figure 1 As shown, an impact-resistant aircraft structure includes an aircraft structure body 10, a first-stage shock absorber 20, a second-stage shock absorber 30, a drive component, a positioning boss, and a gasket 80.

[0058] Multiple internal components 40 are installed inside the spacecraft structural body 10. Multiple first-stage vibration dampers 20 are mounted at different locations on the spacecraft structural body 10 and are configured to reduce the impact overload force on the spacecraft structural body through their own compressive deformation. The first-stage vibration dampers 20 are located between the spacecraft structural body 10 and the drive components. The first-stage vibration dampers 20 provide primary vibration damping for the spacecraft structural body 10.

[0059] Multiple secondary vibration dampers 30 are installed between the intra-satellite unit 40 and the spacecraft structure body 10 to dampen the intra-satellite unit 40.

[0060] The drive component is in close contact with the first-stage damper and is configured to withstand and transfer the impact overload force to the first-stage damper. Close contact between the drive component and the first-stage damper means there are no gaps between their contact surfaces.

[0061] The positioning boss provides radial positioning for the drive component and cooperates with the first-stage damper 20 to fix the drive component. The positioning boss includes a first positioning boss 60 and a second positioning boss 70. The drive component includes a first drive component 51 and a second drive component 52.

[0062] In one embodiment of the present invention, the onboard unit 40 includes a data processing system unit and / or a thermal control system unit and / or a communication system unit and / or an attitude control system unit and / or a power supply system unit.

[0063] The first-stage vibration damper 20 is a structural-level vibration damper that reduces vibration on the main body of the spacecraft structure 10. The second-stage vibration damper 30 is a single-unit-level vibration damper that reduces vibration on the single unit 40 inside the satellite.

[0064] In one embodiment of the present invention, the aircraft structure body 10 includes an aircraft body 101, a first vibration damping mounting structure 102, a second vibration damping mounting structure 103, and a front section structure 104. The aircraft structure body 10 is a one-piece molded structure.

[0065] In one embodiment of the present invention, the front section structure 104 is located at the front end of the aircraft body 101, the first vibration damping mounting structure 102 and the second vibration damping mounting structure 103 are respectively located at both ends of the aircraft body, and wherein the first vibration damping mounting structure 102 is located between the front section structure 104 and the aircraft body 101.

[0066] In one embodiment of the present invention, the first vibration damping mounting structure 102 and the second vibration damping mounting structure 103 are annular, and the diameters of the first vibration damping mounting structure 102 and the second vibration damping mounting structure 103 are larger than the diameter of the aircraft body 101.

[0067] The first vibration damping mounting structure is configured to provide a mounting position for the first-stage vibration damper 20 and prevent the first-stage vibration damper 20 from sliding when subjected to impact overload force.

[0068] In one embodiment of the present invention, the first vibration damping mounting structure 102 includes a first main body portion 1021 and a first blocking portion 1022 and a first mounting portion 1023 located at both ends of the first main body portion 1021 and protruding outward relative to the first main body portion 1021. Here, "outward" refers to the direction away from the aircraft structure body.

[0069] The first main body 1021 provides an installation position for the first-stage vibration damper 20, which is sleeved on the first main body 1021. The first blocking part 1022 is used to prevent the first-stage vibration damper 20 from sliding when subjected to impact overload force. The first mounting part 1023 is used to install the first positioning boss 60, which can be installed on the first mounting part 1023 by bolts.

[0070] In one embodiment of the present invention, the first blocking portion 1022 is located away from the aircraft body 101, and the first mounting portion 1023 is located adjacent to the aircraft body 101. The outer wall of the first main body portion 1021 is recessed inward relative to the first blocking portion 1022 and the first mounting portion 1023. Here, "inward" means in the direction toward the central axis of the aircraft structure body.

[0071] In one embodiment of the present invention, the second vibration damping mounting structure 103 includes a second main body portion 1031 and a second blocking portion 1032. The second blocking portion 1032 protrudes outward relative to the second main body portion 1031 and is adjacent to the aircraft body 101. The second main body portion 1031 provides a mounting position for the first-stage vibration damper 20, which is sleeved on the second main body portion 1031. The second blocking portion 1032 is used to prevent the first-stage vibration damper 20 from sliding when subjected to impact overload force.

[0072] In one embodiment of the present invention, two first-stage vibration dampers 20 are respectively mounted on a first vibration damping mounting structure 102 and a second vibration damping mounting structure 103. The two (or more) first-stage vibration dampers are respectively placed on the exterior of both ends of the aircraft body, in parallel connection. In another embodiment of the present invention, the first-stage vibration damper 20 is a structural-level annular vibration damper, a circumferentially closed structure. The two first-stage vibration dampers 20 are sleeved on the first vibration damping mounting structure 102 and the second vibration damping mounting structure 103, and the size of the vibration damper can be adjusted axially as needed.

[0073] In one embodiment of the present invention, the first-stage shock absorber 20 is made of a high-energy-density metal material and is integrally molded and fitted onto the outside of the aircraft structure body 10. It achieves a large degree of impact attenuation by deforming itself under compression.

[0074] In one embodiment of the present invention, such as Figure 2 and 3 As shown, the overall structure of the first-stage vibration damper 20 resembles a bellows configuration. The first-stage vibration damper 20 is made by heating and pressing a thin, uniform-wall-thickness, ultra-elastic alloy sheet into a corrugated shape using a mold, then winding it into a ring, and finally welding it into a single piece. The material of the first-stage vibration damper 20 is ultra-elastic titanium alloy.

[0075] After the aircraft structure body 10 is manufactured, the first-stage vibration damper 20 can be formed on the first vibration damping mounting structure and the second vibration damping mounting structure by an integral molding method. The first-stage vibration damper 20 is fixedly connected to the aircraft main structure 10.

[0076] In one embodiment of the present invention, the first end 201 and the second end 202 of the first-stage damper 20 are planar, and the first end 201 and the second end 202 are opposite to each other.

[0077] In one embodiment of the invention, the first-stage damper 20 includes N / 2 corrugated elements 203, where N ≥ 3. The corrugated elements have a flattened V-shaped cross-section.

[0078] All edges of the first-stage vibration damper 20 are designed with rounded corners to prevent damage caused by localized stress during operation. Specifically, the abrupt structural changes in the first-stage vibration damper 20 are designed with rounded corners. Compared to sharp corners, rounded corners allow for smoother force transmission and more uniform stress distribution, preventing stress concentration in a small area. Furthermore, the edges of the corrugated units 203 and the corrugated units 203 between each other also employ rounded corner designs.

[0079] The first-stage vibration damper 20 is installed on the outside of the aircraft structure body 10. Its basic structure is similar to a bellows configuration, which facilitates the aircraft's attitude and trajectory adjustment. At the same time, the layout at both ends can reduce the local stress of the aircraft structure body.

[0080] In one embodiment of the present invention, such as Figure 1 As shown, the first driving component 51 and the second driving component 52 are generally in the shape of a ring. The first driving component 51 surrounds the first vibration damping mounting structure 102 and the first-stage vibration damper 20 installed on the first vibration damping mounting structure 102. The second driving component 52 surrounds the second vibration damping mounting structure 103 and the first-stage vibration damper 20 installed on the second vibration damping mounting structure 103.

[0081] In one embodiment of the present invention, such as Figure 1 As shown, the inner sides of the first drive component 51 and the second drive component 52 both have bearing bosses 53. The first-stage damper 20 is located on the bearing bosses 53. The bearing bosses 53 are in close contact with the first-stage damper 20, which facilitates the first drive component 51 to transmit the impact overload force to the first-stage damper 20.

[0082] In one embodiment of the present invention, the structure and dimensions of the first driving component 51 and the second driving component 52 may not be exactly the same. For example, the bearing bosses 53 of the first driving component 51 and the second driving component 52 may be located at different positions of the first driving component 51 and the second driving component 52. The heights of the first driving component 51 and the second driving component 52 may be different; for example, the height of the second driving component 52 may be greater than that of the first driving component 51.

[0083] Multiple first-stage vibration dampers 20 are installed inside the first drive component 51 and the second drive component 52 and outside the aircraft structure body 10, and are axially positioned by the bearing boss 53 and the blocking parts (first blocking part 1022 and second blocking part 1032) on the aircraft structure body 10.

[0084] In one embodiment of the present invention, such as Figure 1 As shown, the first positioning boss 60 is sleeved on the first vibration damping mounting structure 102 and fixedly connected to the first mounting part 1023 by bolts. The first positioning boss 60 is generally annular, and its inner side contacts the first main body part 1021 of the first vibration damping mounting structure 102. The first positioning boss 60 includes a first radial positioning part 61 and a first bearing part 62, both with the same inner diameter, and the outer diameter of the first bearing part 62 is larger than the outer diameter of the first radial positioning part 61. The first positioning boss 60 is generally annular.

[0085] like Figure 1 As shown, the first driving component 51 is sleeved on the first positioning boss 60. The bearing boss 53 of the first driving component 51 is located on the first bearing portion 62 of the first positioning boss 60, and the bearing boss 53 of the first driving component 51 is located below the first-stage vibration damper 20. The inner surface of the bearing boss 53 of the first driving component 51 is in contact with the outer surface of the first radial positioning portion 61 of the first positioning boss 60. The first positioning boss 60 cooperates with the first-stage vibration damper 20 to fix the first driving component 51, without the need for fasteners, only close contact. The first positioning boss 60 provides radial positioning for the first driving component, ensuring that the first driving component 51 will not move downwards or left or right.

[0086] In one embodiment of the present invention, such as Figure 1 As shown, the second positioning boss 70 is fastened to the end of the second vibration damping mounting structure 103 by bolts. The second positioning boss 70 is generally ring-shaped.

[0087] The second positioning boss 70 includes a second radial positioning part 71 and a second bearing part 72. The outer diameter of the second bearing part 72 is larger than the outer diameter of the second radial positioning part 71, that is, the second bearing part 72 protrudes outward relative to the second radial positioning part 71.

[0088] In one embodiment of the present invention, such as Figure 1 As shown, the gasket 80 is annular, fitted onto the second positioning boss 70, and located on the upper surface of the second bearing portion 72. Furthermore, the gasket 80 is located between the second bearing portion 72 and the bearing boss 53 of the second driving component 52.

[0089] like Figure 1As shown, the second driving component 52 is sleeved on the second positioning boss 70. The bearing boss 53 of the second driving component 52 is located on the second bearing portion 72 of the second positioning boss 70, and the gasket 80 is located between the second bearing portion 72 and the bearing boss 53 of the second driving component 52. The bearing boss 53 of the second driving component 52 is located below the first-stage vibration damper 20, and the inner surface of the bearing boss 53 of the second driving component 52 contacts the outer surface of the second radial positioning portion 71 of the second positioning boss 70. The second positioning boss 70 cooperates with the first-stage vibration damper 20 to fix the second driving component 52, without the need for fasteners, only close contact. The second positioning boss 70 provides radial positioning for the first driving component, ensuring that the second driving component 52 will not move downwards or left or right.

[0090] The first-stage vibration damper 20, installed on the first vibration damping mounting structure 102, connects the first vibration damping mounting structure 102 and the first drive component 51. The first-stage vibration damper 20, installed on the second vibration damping mounting structure 103, connects the second vibration damping mounting structure 103 and the second drive component 52. The first-stage vibration damper 20 adopts an integrated structural and functional design with high body rigidity. It can serve as a connecting component between the drive component and the aircraft structural body 10. When subjected to high-level overloads, it undergoes compressive deformation to achieve a significant vibration damping effect.

[0091] In one embodiment of the present invention, the interior of the spacecraft body 101 is hollow, allowing for the installation of internal units 40. Multiple internal units 40 are installed on the inner wall of the spacecraft body 101, and a second-stage vibration damper 30 is disposed between the internal unit 40 and the inner wall of the spacecraft body 101. Furthermore, the internal units 40 are fastened to the spacecraft body 101 with bolts, and each bolt is fitted with a second-stage vibration damper 30, which is located between the internal unit 40 and the spacecraft body 101.

[0092] The second-stage vibration damper 30 is a single-unit point-type vibration damper, which can be a pad-type ring vibration damper. The material, size, specifications, and performance of the vibration damper are selected according to the vibration damping requirements of different single units. For example, the second-stage vibration damper 30 can be a silicon-based (Si) rubber vibration damper or a metal-rubber vibration damper (a metal vibration damper made of high-elasticity fine metal wires such as NiTi alloy or aluminum alloy), which are commonly used in aerospace applications. This type of product is a mature racking product, available in a series of specifications, and can be selected according to the actual conditions of a single unit location.

[0093] In one embodiment of the present invention, such as Figure 1As shown, the front structure 104 is a ring-shaped structure with a hollow interior, allowing the installation of an intra-satellite unit. A second-stage vibration damper can also be installed between the intra-satellite unit and the front structure 104. The front structure 104 includes multiple support columns 1041 and a ring-shaped end 1042. The two ends of the multiple support columns 1041 are respectively connected to the ring-shaped end 1042 and the first vibration damping mounting structure 102.

[0094] The principle of secondary vibration reduction is explained below with reference to the attached diagram.

[0095] like Figure 4 As shown, this invention utilizes a structural-level annular vibration damper (first-stage vibration damper 20) as a primary vibration damping structure, participating in the transmission chain of the main load-bearing structure to resist instantaneous large impacts. It achieves a large degree of impact attenuation through its own compression deformation, realizing primary vibration damping and protecting the spacecraft structure and internal single units. Then, the impact on the single unit inside the spacecraft is further reduced by the single-unit-level vibration damper.

[0096] The spacecraft withstands high-level impact loads through its drive components. Kinetic energy is transferred via structural-level ring dampers, enabling the spacecraft structure 10 and the internal unit 40 to achieve initial velocity. Simultaneously, the internal unit 40 undergoes secondary vibration reduction via unit-level point-position dampers, further reducing the mechanical input transmitted to the unit. Using two (or more) sets of parallel structural-level ring dampers as the primary vibration reduction of the system achieves at least 50% impact attenuation. Using unit-level point-position dampers as the secondary vibration reduction reduces the remaining vibration magnitude to below 2000g, effectively lowering the impact magnitude experienced by each unit.

[0097] Current vibration reduction methods only use single-stage vibration reduction, which cannot achieve near-10,000 g (100,000 m / s) levels in the time domain under limited space and mass constraints. 2 The high attenuation ratio of impacts at the (level) is achieved. The two-stage composite vibration reduction scheme of this invention, at both the structural and single-unit levels, can achieve near-10,000 g (100,000 m / s) time-domain vibrations under limited space and mass constraints. 2 The high attenuation ratio of the shock absorber (at level 1) protects the normal operation of the aircraft. Compared to conventional single-unit-level shock absorption, the aircraft structure of this invention adopts a two-stage shock absorption system, which can cope with a large number of shock overload forces.

[0098] While some embodiments of the present invention have been described in this application, those skilled in the art will understand that these embodiments are merely illustrative. Numerous variations, alternatives, and improvements will arise in those skilled in the art under the teachings of this invention without departing from its scope. The appended claims are intended to define the scope of the invention and thereby cover methods and structures within the scope of the claims themselves and their equivalents.

Claims

1. A high-impact aircraft structure with secondary vibration reduction, characterized in that, include: The spacecraft's main structure contains multiple internal satellite units; Multiple primary shock absorbers are fitted onto the aircraft structure and configured to reduce the impact overload force on the aircraft structure through their own compressive deformation. Multiple secondary shock absorbers are disposed between the multiple intra-satellite units and the spacecraft structure to further reduce the impact overload force on the intra-satellite units; as well as A drive component, which is in close contact with the first-stage damper and is configured to withstand impact overload forces and transmit the impact overload forces to the first-stage damper; The aircraft structure body is a one-piece molded structure, and the aircraft structure body includes: The main body of the spacecraft is hollow inside, and multiple onboard units are installed inside the main body of the spacecraft. A front section structure, located at the front end of the main body of the aircraft, and the front section structure is hollow inside; A first vibration damping mounting structure, located between the front structure and the aircraft body, is configured to provide a mounting position for the first-stage vibration damper and prevent the first-stage vibration damper from sliding under impact overload forces; and The second vibration damping mounting structure is located at the end of the aircraft body and is configured to provide a mounting position for the first-stage vibration damper and prevent the first-stage vibration damper from sliding when subjected to impact overload force; The first vibration damping mounting structure includes: The first main body is configured to provide a mounting position for the first stage vibration damper; A first blocking portion, located at one end of the first main body portion, protrudes outward relative to the first main body portion and is configured to prevent the first-stage shock absorber from sliding under impact overload force; and A first mounting portion is located at the other end of the first main body portion and is connected to the drive component; and / or The second vibration damping mounting structure includes: The second main body is configured to provide a mounting position for the first-stage vibration damper; and The second blocking part, located at one end of the second main body, protrudes outward relative to the second main body and is configured to prevent the first-stage shock absorber from sliding when subjected to impact overload force.

2. The impact-resistant aircraft structure with secondary vibration reduction according to claim 1, characterized in that, The plurality of first-stage vibration damper sleeves are annular vibration dampers, which are integrally molded and fitted onto the first vibration damping mounting structure and the second vibration damping mounting structure, including: A metal plate of uniform wall thickness is heated and pressed into a corrugated shape using a mold, then wound into a ring, and finally welded together to form a single unit, thus obtaining the first-stage vibration damper.

3. The impact-resistant aircraft structure with secondary vibration reduction according to claim 2, characterized in that, The first-stage vibration damper comprises N / 2 corrugated elements, where N ≥ 3; and / or All edges of the first-stage vibration damper are rounded.

4. The impact-resistant aircraft structure with secondary vibration reduction according to claim 1, characterized in that, The driving component includes a first driving component and a second driving component, wherein the first driving component surrounds the first vibration damping mounting structure and a first-stage vibration damper mounted on the first vibration damping mounting structure, and the second driving component surrounds the second vibration damping mounting structure and a first-stage vibration damper mounted on the second vibration damping mounting structure. Both the first driving component and the second driving component have a bearing boss on their inner sides, and the bearing boss is in close contact with the first-stage vibration damper.

5. The impact-resistant aircraft structure with secondary vibration reduction according to claim 4, characterized in that, Also includes: A positioning boss is configured to provide radial positioning for the drive component and to cooperate with the first-stage damper to fix the drive component.

6. The impact-resistant aircraft structure with secondary vibration reduction according to claim 5, characterized in that, The positioning boss includes: A first positioning boss is sleeved on the first vibration damping mounting structure and fixedly connected to the first mounting part. The first positioning boss includes a first radial positioning part and a first bearing part, wherein the outer diameter of the first bearing part is larger than the outer diameter of the first radial positioning part; and / or The second positioning boss is installed at the end of the second vibration damping mounting structure. The second positioning boss includes a second radial positioning part and a second bearing part, wherein the outer diameter of the second bearing part is larger than the outer diameter of the second radial positioning part.

7. The impact-resistant aircraft structure with secondary vibration reduction according to claim 6, characterized in that, The first driving component is sleeved on the first positioning boss. The bearing boss of the first driving component is located on the first bearing part of the first positioning boss and below the first stage vibration damper. The inner surface of the bearing boss of the first driving component is in contact with the outer surface of the first radial positioning part of the first positioning boss.

8. The impact-resistant aircraft structure with secondary vibration reduction according to claim 6, characterized in that, It also includes a gasket, which is fitted onto the second positioning boss and located between the second bearing portion and the bearing boss of the second driving component.

9. The impact-resistant aircraft structure with secondary vibration reduction according to claim 8, characterized in that, The second driving component is sleeved on the second positioning boss. The bearing boss of the second driving component is located above the pad and below the first-stage damper. The inner surface of the bearing boss of the second driving component is in contact with the outer surface of the second radial positioning part of the second positioning boss.

10. The high-impact aircraft structure with secondary vibration reduction according to claim 1, characterized in that, The multiple intra-satellite units are fastened to the inner wall of the spacecraft body by bolts, and each bolt is fitted with a second-stage vibration damper, wherein the second-stage vibration damper is a silicon-based rubber vibration damper or a metal-rubber vibration damper.

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