Instantaneous large overload resistant spacecraft structure

Abstract

The invention relates to an instantaneous large overload resistant spacecraft structure, which comprises: a main body structure, which comprises a plurality of bearing columns, the outer surfaces of which define a revolution body and are configured to bear an overload impact force; the support beams are arranged between every two adjacent force bearing columns and are configured to provide mounting points for the single aircraft; the end connecting structures are located at the two ends of the main body structure and provided with first threads. The spacecraft structure adopts a gyration body configuration, the stability and strength of the structure are effectively improved, the transverse rigidity of the structure is increased through the gyration body structure, the pointing precision of the spacecraft is effectively improved, the critical yield load coefficient of the structure is increased, and the problem that the structure is damaged due to configuration instability when subjected to large overload impact is solved; and uniform stress of the structure is ensured when the structure bears instantaneous overload, and damage caused by stress concentration is avoided.

Description

Technical Field

[0001] The present invention relates to the field of aircraft technology, and in particular to a space aircraft structure capable of resisting instantaneous large overload. Background Art

[0002] At present, the demand for anti-overload of the structure of space-based systems is mainly due to the complex mechanical environment of the active section of the launch vehicle. The structure of the satellite can usually cope with 9-12g static overload, and generally there is no need to make adaptive design for the overall configuration of the satellite. However, when facing a force close to 10,000g (100,000 m / s 2 When the working environment is subjected to large instantaneous (time domain) overload (level 1), the traditional satellite structure will no longer meet the strength requirements and conventional structural reinforcement measures will also fail.

[0003] The existing spacecraft structure has the following problems:

[0004] The structural strength and rigidity of the spacecraft of the existing space-based system cannot cope with high-level overload conditions, and the configuration stability cannot cope with large overload conditions of nearly 10,000 g in the time domain.

[0005] The structural connection method of traditional aircraft cannot cope with the high-level instantaneous overload environment of nearly 10,000 g in the time domain, and will disintegrate and be destroyed.

[0006] Traditional large overload-resistant structures are usually continuous solid structures with too much mass and are not suitable for space environments. Conventional frame structures cannot cope with large-scale impact overloads. Summary of the invention

[0007] In order to solve at least part of the above problems in the prior art, the task of the present invention is to provide a spacecraft structure that resists instantaneous large overload, comprising:

[0008] The main structure includes:

[0009] A plurality of load-bearing columns, the outer surfaces of which form a body of revolution and are configured to withstand overload impact force; and

[0010] A plurality of support beams, which are disposed between two adjacent load-bearing columns and are configured to provide mounting points for a single aircraft; and

[0011] A plurality of end connection structures are located at two ends of the main structure and have a first thread.

[0012] Furthermore, the cross section of the load-bearing column is a special-shaped symmetrical cross section;

[0013] The load-bearing column has two mutually perpendicular side surfaces and an arc-shaped outer surface connecting the two side surfaces; and

[0014] The internal space surrounded by the multiple load-bearing columns is a cube or a cuboid.

[0015] Furthermore, the cross section of the support beam is rectangular, and the length-to-width ratio of the cross section of the support beam is 3:1-5:1.

[0016] Furthermore, the end connection structure comprises:

[0017] A first annular structure connected to the plurality of load-bearing columns;

[0018] A second annular structure having a first thread on its inner side;

[0019] A plurality of first connecting columns connect the first annular structure and the second annular structure and are configured to withstand overload impact force.

[0020] Furthermore, the first connecting column has an arc-shaped outer surface and an inner surface, and the area of ​​the inner surface is smaller than the area of ​​the outer surface.

[0021] Furthermore, it also includes:

[0022] The front section structure includes:

[0023] A third annular structure, which is threadedly connected to the end connection structure located at the first end of the main structure, the third annular structure has a first connection portion, and the outer surface of the first connection portion is provided with a second thread;

[0024] a fourth annular structure, which is away from the main structure; and

[0025] a plurality of second connecting posts connecting the third annular structure and the fourth annular structure; and

[0026] The rear section structure is threadedly connected to the end connection structure located at the second end of the main structure. The rear section structure has a second connection portion, and the outer surface of the second connection portion is provided with a third thread.

[0027] Furthermore, the width of the thread teeth of the first thread, the second thread and the third thread is 1-4 mm, and the length is more than 10 mm.

[0028] Furthermore, the material of the main structure, the end connection structure, the front section structure and the rear section structure is Ti6Al4V titanium alloy.

[0029] Furthermore, the main body structure and the end connection structure are formed into an integrated structure;

[0030] The front section structure is formed into an integrated structure; and

[0031] The rear section structure is formed into an integrated structure.

[0032] Furthermore, the angle between the load-bearing column and the support beam is a rounded angle;

[0033] The angle between the first annular structure and the outer surface of the first connecting column in the end connecting structure is a rounded angle;

[0034] In the front section structure, multiple included angles between the second connecting column and the third annular structure and the fourth annular structure are chamfered.

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

[0036] (1) The spacecraft structure adopts a gyroscopic configuration, which effectively improves the stability and strength of the structure. The gyroscopic structure increases the lateral stiffness of the structure and effectively improves the pointing accuracy of the spacecraft.

[0037] (2) By designing the gyroscopic configuration of the aircraft body, the critical yield load coefficient of the structure is increased, which solves the problem of structural damage caused by configuration instability when subjected to large overload impact, and ensures that the structure is uniformly stressed when subjected to instantaneous overload, avoiding stress concentration and damage.

[0038] (3) The cross section of the main load-bearing structure (load-bearing column) of the aircraft adopts a special-shaped symmetrical cross-section design, which can maximize the overload resistance of the structure while reducing the mass of the structure itself. The support beam adopts a high aspect ratio cross-section beam structure, which can increase the bending and torsional strength of the support along the overload application direction, effectively reduce the overall mass of the structure, and provide high-strength connection support for the single-machine installation point.

[0039] (4) Through the hollow cube space envelope design inside the main structure of the aircraft, space is reserved for installing a single aircraft, solving the problem of low internal space utilization of the gyroscopic configuration.

[0040] (5) The use of segmented splicing structure in spacecraft can solve the assembly process problem. The various segments of the spacecraft are screwed together through coarse thread design to achieve high-strength connection, which resolves the contradiction between the overall strength requirement of the spacecraft and the internal single-machine installation process, and solves the strength problem of traditional structural connection methods that cannot withstand large overload impacts. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] In order to further illustrate the above and other advantages and features of various embodiments of the present invention, a more specific description of various embodiments of the present invention will be presented with reference to the accompanying drawings. It will be understood that these drawings only depict typical embodiments of the present invention and are therefore not to be considered as limiting the scope thereof. In the accompanying drawings, for clarity, identical or corresponding parts will be represented by identical or similar reference numerals.

[0042] Figure 1A longitudinal cross-sectional schematic diagram of a spacecraft structure resistant to instantaneous large overload according to an embodiment of the present invention is shown;

[0043] Figure 2 A perspective view showing a spacecraft structure for resisting instantaneous large overload according to an embodiment of the present invention;

[0044] Figure 3 A top view showing a space vehicle structure according to one embodiment of the present invention; and

[0045] Figure 4 A schematic diagram showing a cross section of a mid-section structure according to an embodiment of the present invention. DETAILED DESCRIPTION

[0046] It should be noted that various components in the various drawings may be shown exaggeratedly for illustrative purposes and are not necessarily correct to scale.

[0047] In the present invention, each embodiment is only intended to illustrate the aspects of the present invention and should not be construed as limiting.

[0048] In the present invention, unless otherwise specified, the quantifiers "a", "an" and "an" do not exclude the presence of a plurality of elements.

[0049] It should also be pointed out that in the embodiments of the present invention, for the sake of clarity and simplicity, only a portion of the parts or components may be shown, but a person of ordinary skill in the art will understand that under the teachings of the present invention, required parts or components may be added as needed in specific scenarios.

[0050] It should also be pointed out that within the scope of the present invention, the terms "same", "equal", "equal to" and the like do not mean that the two values ​​are absolutely equal, but allow a certain reasonable error, that is, the terms also cover "substantially the same", "substantially equal", and "substantially equal to".

[0051] It should also be pointed out that in the description of the present invention, the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside" and the like indicate positions or positional relationships based on the positions or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than explicitly or implicitly indicating that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as limiting the present invention. In addition, the terms "first" and "second" are used for descriptive purposes only and cannot be understood as explicitly or implicitly indicating relative importance.

[0052] In addition, the embodiments of the present invention describe the process steps in a specific order, but this is only for the convenience of distinguishing the steps, and does not limit the sequence of the steps. In different embodiments of the present invention, the sequence of the steps can be adjusted according to the adjustment of the process.

[0053] In the present invention, the term "configuration" refers to the setting of the shape, structure, material and / or function of the target object to achieve the desired technical effect, wherein "configuration" includes a variety of alternative technical means for achieving the technical effect, which become obvious under the teaching of the present invention.

[0054] Figure 1 A longitudinal cross-sectional schematic diagram of a spacecraft structure resistant to instantaneous large overload according to an embodiment of the present invention is shown; Figure 2 A perspective view showing a spacecraft structure for resisting instantaneous large overload according to an embodiment of the present invention; Figure 3 A top view of a space vehicle structure according to an embodiment of the present invention is shown; Figure 4 A schematic diagram showing a cross section of a mid-section structure according to an embodiment of the present invention.

[0055] like Figure 1 As shown, a spacecraft structure resistant to transient large overload comprises a front section structure 10, a spacecraft body 20 and a rear section structure 30. The spacecraft body 20 is connected to the front section structure 10 and the rear section structure 30 by threading.

[0056] In one embodiment of the present invention, the front section structure 10, the aircraft body 20 and the rear section structure 30 are all made of high-strength Ti 6Al4V titanium alloy as the body material, and high-precision machining technology is used to achieve integrated structural molding. The aircraft structure is completed by integrated machining of high-strength titanium alloy, with high overall strength and high rigidity, and can effectively bear overload impacts close to 10,000 g in the time domain without structural damage.

[0057] In one embodiment of the present invention, Figure 2 and 3 As shown, the exterior of the aircraft body 20 is a gyroscopic configuration (rotating body), the overall shape is cylindrical, the main internal area is a space cube, and the internal space of the aircraft body 20 can be installed with a single machine. The main internal area can also be a cuboid. The single machine can include a data processing system single machine and / or a thermal control system single machine and / or a communication system single machine and / or an attitude control system single machine and / or a power system single machine.

[0058] The aircraft body 20 includes a main structure 21 and end connection structures 22 located at both ends of the main structure. The main structure 21 and the end connection structures 22 are formed as an integrated structure.

[0059] The main structure 21 includes a plurality of bearing columns 211 and a plurality of support beams 212. A plurality of support beams 212 are connected between two adjacent bearing columns 211. The support beams 212 serve as connecting brackets of the bearing columns 211, and the brackets also serve as installation points of the single machine.

[0060] Furthermore, the main structure 21 includes four bearing columns 211, two support beams 212 are connected between two adjacent bearing columns 211, and the main structure 21 includes eight support beams 212. The aircraft body is composed of four groups (or more groups) of main bearing columns to form a core bearing structure for bearing overload impact force.

[0061] In one embodiment of the present invention, two ends of the plurality of load-bearing columns 211 are connected to two end connection structures 22 .

[0062] In one embodiment of the present invention, Figure 4 As shown, the cross section of the load-bearing column 211 is a special-shaped symmetrical cross section, and the load-bearing column 211 has two mutually perpendicular side surfaces and an arc-shaped outer surface connecting the two side surfaces. The two side surfaces of the load-bearing column 211 are mutually perpendicular, so that the main structure 21 has a symmetrical configuration. The special-shaped symmetrical cross section is a fan-shaped cross section. Figure 2 As shown, if the arc-shaped outer surfaces of multiple load-bearing columns 211 are connected, they can form a spiral outer contour.

[0063] The load-bearing column 211 adopts a special-shaped symmetrical cross-section design mainly to meet two requirements and minimize the mass. On the one hand, it is necessary to ensure that the outer contour of the main load-bearing structure of the structure is a gyroscope configuration to effectively increase the shortest distance of mechanical transmission. On the other hand, it must also meet the layout of the installation support beam 212.

[0064] In one embodiment of the present invention, Figure 1 As shown, the cross section of the support beam 212 is rectangular, and the long side of the cross section of the support beam 212 is parallel to the axis of the gyroscope structure.

[0065] Furthermore, the aspect ratio of the cross section of the support beam 212 is 3:1-5:1. The support beam adopts a high aspect ratio cross-sectional beam structure, that is, the dimension along the bending direction is larger than the dimension perpendicular to the bending direction, and the bending resistance is better under equal cross-sectional area.

[0066] In one embodiment of the present invention, Figure 1 and 2 As shown, the end connection structure 22 includes a first annular structure 221 , a second annular structure 222 , and a plurality of first connection columns 223 connecting the first annular structure 221 and the second annular structure 222 .

[0067] The first annular structure 221 is adjacent to the main structure 21 and connected to a plurality of load-bearing columns 211. The second annular structure 222 is away from the main structure 21.

[0068] In one embodiment of the present invention, the inner side of the second annular structure 222 has a first thread 224. The thread teeth of the first thread 224 have a width of 1-4 mm and a length of more than 10 mm. The shape of the thread teeth can be rectangular or trapezoidal.

[0069] The outer surface and the inner surface of the first connecting column 223 are arc-shaped surfaces, and the area of ​​the inner surface is smaller than the area of ​​the outer surface. Further, the cross section of the first connecting column 223 is trapezoidal.

[0070] In one embodiment of the present invention, Figure 1 As shown, the angles in the aircraft body 20 are mostly rounded. For example, the angle between the load-bearing column 211 and the support beam 212 is rounded. The angle between the first annular structure 221 in the end connection structure 22 and the outer surface of the first connection column 223 is rounded.

[0071] In one embodiment of the present invention, Figure 1 As shown, the front section structure 10 includes a third annular structure 11, a fourth annular structure 12, and a plurality of second connecting columns 13 connecting the third annular structure 11 and the fourth annular structure 12. The front section structure 10 is threadedly connected to the aircraft body 20.

[0072] The front section structure 10 is threadedly connected to the end connection structure 22 located at the first end of the main structure 21 .

[0073] The third annular structure 11 has a first connecting portion 14, and the outer surface of the first connecting portion 14 is provided with a second thread 15. The thread teeth of the second thread 15 have a width of 1-4 mm and a length of more than 10 mm. The shape of the thread teeth can be rectangular or trapezoidal. The first connecting portion 14 is used to connect the front section structure 10 to the aircraft body 20.

[0074] The outer surface and the inner surface of the second connecting column 13 are both arc-shaped surfaces. The second connecting column 13 constitutes the core load-bearing structure of the front section structure 10 and is used to withstand overload impact force.

[0075] A plurality of included angles between the second connecting column 13 and the third annular structure 11 and the fourth annular structure 12 are rounded.

[0076] The outer diameter of the fourth annular structure 12 is greater than the outer diameter of the third annular structure 11 .

[0077] In one embodiment of the present invention, Figure 1As shown, the rear section structure 30 is in a flat ring shape as a whole. The rear section structure 30 has a second connecting portion 31, and the outer surface of the second connecting portion 31 is provided with a third thread 32. The thread teeth of the third thread 32 have a width of 1-4 mm and a length of more than 10 mm. The shape of the thread teeth can be rectangular or trapezoidal. The second connecting portion 31 is used to connect the rear section structure 30 to the aircraft body 20. The rear section structure 30 is threadedly connected to the end connection structure 22 located at the second end of the main structure 21.

[0078] The threads on the inner side of the second annular structure 222 of the aircraft body 20, the threads of the first connecting portion 14 of the front section structure 10, and the threads of the second connecting portion 31 of the rear section structure 30 are all designed with coarse pitch threads.

[0079] The aircraft structure of the present invention has the following advantages:

[0080] The space vehicle structure adopts a gyroscopic configuration. The cross-sectional profile of the gyroscopic configuration is circular, and the force on the main load-bearing structure is more uniform, which can effectively improve the stability and strength of the structure. The gyroscopic structure increases the lateral stiffness of the structure and effectively improves the pointing accuracy of the vehicle.

[0081] In addition, the cross-sectional profile of the body of revolution configuration is a circle, and the circular area is the largest when the circumference is equal, which can more effectively utilize limited materials and provide the largest single-machine layout space; in addition, the aircraft structure adopts integrated machining, and the body of revolution configuration is easier to implement in processing.

[0082] By designing the gyroscopic configuration of the outside of the spacecraft body, the critical yield load coefficient of the structure is increased, the problem of structural damage due to configuration instability when subjected to a large overload impact is solved, and uniform force is ensured on the structure when subjected to instantaneous overload, avoiding damage caused by stress concentration.

[0083] The cross section of the aircraft's main load-bearing structure (load-bearing column) adopts a special-shaped symmetrical cross-section design, which can maximize the structure's anti-overload performance while reducing the mass of the structure itself. The support beam adopts a high aspect ratio cross-section beam structure, with the dimension along the bending direction larger than the dimension perpendicular to the bending direction. This structure can increase the support's bending and torsional strength along the overload application direction, effectively reducing the overall mass of the structure and providing high-strength connection support for single-machine installation points.

[0084] Through the hollow cube space envelope design inside the main structure of the aircraft, space is reserved for installing a single machine, thus solving the problem of low internal space utilization of the gyroscopic configuration.

[0085] The aircraft structure can adopt a segmented splicing structure (two or three sections) to solve the assembly process problem (it can be not segmented when there is no conflict with the assembly process). The various sections of the aircraft structure are tightened together through a coarse thread design to achieve a high-strength connection of the main structure, resolving the contradiction between the overall strength requirements of the aircraft and the internal single-machine installation process, and solving the strength problem that the traditional structural connection method cannot withstand large overload impacts.

[0086] Although some embodiments of the present invention have been described in this application document, it will be appreciated by those skilled in the art that these embodiments are merely shown as examples. Those skilled in the art may conceive of numerous variations, alternatives, and improvements under the teachings of the present invention without departing from the scope of the present invention. The appended claims are intended to define the scope of the present invention and to cover methods and structures within the scope of these claims themselves and their equivalents.

Claims

1. A spacecraft structure capable of resisting instantaneous large overload, characterized in that: include: The main structure includes: A plurality of load-bearing columns, the outer surfaces of which form a body of revolution and are configured to withstand overload impact force; and A plurality of support beams, which are disposed between two adjacent load-bearing columns and are configured to provide mounting points for a single aircraft; and A plurality of end connection structures are located at two ends of the main structure and have a first thread.

2. The spacecraft structure for resisting instantaneous large overload according to claim 1, characterized in that: The cross section of the load-bearing column is a special-shaped symmetrical cross section; The load-bearing column has two mutually perpendicular side surfaces and an arc-shaped outer surface connecting the two side surfaces; and The internal space surrounded by the multiple load-bearing columns is a cube or a cuboid.

3. The spacecraft structure for resisting instantaneous large overload according to claim 2, characterized in that: The cross section of the support beam is rectangular, and the length-to-width ratio of the cross section of the support beam is 3:1-5:

1.

4. The spacecraft structure for resisting instantaneous large overload according to claim 1, characterized in that: The end connection structure comprises: A first annular structure connected to the plurality of load-bearing columns; A second annular structure having a first thread on its inner side; A plurality of first connecting columns connect the first annular structure and the second annular structure and are configured to withstand overload impact force.

5. The spacecraft structure for resisting instantaneous large overload according to claim 4, characterized in that: The first connecting column has an arc-shaped outer surface and an inner surface, and the area of ​​the inner surface is smaller than the area of ​​the outer surface.

6. The spacecraft structure for resisting transient large overload according to claim 4, characterized in that: Also includes: The front section structure includes: A third annular structure, which is threadedly connected to the end connection structure located at the first end of the main structure, the third annular structure has a first connection portion, and the outer surface of the first connection portion is provided with a second thread; a fourth annular structure, which is away from the main structure; and a plurality of second connecting posts connecting the third annular structure and the fourth annular structure; and The rear section structure is threadedly connected to the end connection structure located at the second end of the main structure. The rear section structure has a second connection portion, and the outer surface of the second connection portion is provided with a third thread.

7. The spacecraft structure for resisting instantaneous large overload according to claim 6, characterized in that: The thread teeth of the first thread, the second thread and the third thread have a width of 1-4 mm and a length of more than 10 mm.

8. The spacecraft structure for resisting instantaneous large overload according to claim 6, characterized in that: The main structure, the end connection structure, the front section structure and the rear section structure are made of Ti 6Al4V titanium alloy.

9. The spacecraft structure capable of resisting transient large overload according to claim 6, characterized in that: The main body structure and the end connection structure are formed into an integrated structure; The front section structure is formed into an integrated structure; and The rear section structure is formed into an integrated structure.

10. The spacecraft structure capable of resisting instantaneous large overload according to claim 6, characterized in that: The angle between the load-bearing column and the support beam is a rounded angle; The angle between the first annular structure and the outer surface of the first connecting column in the end connecting structure is a rounded angle; In the front section structure, multiple included angles between the second connecting column and the third annular structure and the fourth annular structure are chamfered.

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