Sliding unfolding device for micro-nano satellite

Abstract

The utility model discloses a sliding unfolding device for a micro-nano satellite, and relates to the technical field of satellites. According to the technical key points, the satellite comprises a satellite main body, solar wing plates are hinged to the opposite sides of the satellite main body respectively, and elastic driving pieces are arranged at the joints of the solar wing plates and the satellite main body; a telescopic supporting unit is installed in a shell of the satellite body, the top of the supporting unit is connected with a horizontal linkage rod, and the two ends of the horizontal linkage rod are slidably connected with the shell of the satellite body in the vertical direction. The two ends of the horizontal linkage rod are each provided with a clamping head, and the two solar wing plates are each provided with a clamping groove. An adjusting unit capable of controlling the position of the horizontal linkage rod at the top of the supporting unit in the vertical direction is arranged on the upper side of the middle position of the horizontal linkage rod. According to the unfolding device, the solar wing plate can be rapidly unfolded according to needs, and components in a satellite can be effectively prevented from being damaged by high temperature when the unfolding device is used.

Description

TECHNICAL FIELD

[0001] The application relates to the technical field of satellites, in particular to a sliding unfolding device for micro-nano satellites. BACKGROUND

[0002] Micro-nano satellites are a rapidly developing field of space technology in recent years, mainly referring to small satellites with small mass (usually weighing between one kilogram and one hundred kilograms). They have become an important driver of the commercialization of space technology due to their low cost, rapid development, flexible deployment and other advantages. In particular, the cubic satellites in standardized micro-nano satellites greatly reduce the threshold and cost of research and use through modular design. At present, micro-nano satellites are mainly used in space experiments, earth observation and other projects by universities and research institutions; verification experiments of new sensors and communication technologies; earth observation; low-orbit communication and Internet of Things. Solar wing plates are important devices for micro-nano satellites to obtain energy, and usually adopt a folding structure. After the satellite is launched into orbit, the solar wing unfolding mechanism is used to unfold it. Considering the cost of micro-nano satellites, the unfolding mechanism used by micro-nano satellites is usually simple. A common micro-nano satellite unfolding mechanism generally works as follows: first, the solar wing plate is folded up by a rope. When it is necessary to unfold the solar wing plate, a heater is used to melt the rope, so that the solar wing plate is released. However, in the implementation of this method, the heat generated by the heater is conducted to the components inside the satellite through the connecting piece, thereby damaging the satellite components. CONTENT OF THE INVENTION

[0003] The application provides a sliding unfolding device for micro-nano satellites, which can quickly unfold the solar wing plate as needed, and effectively avoid high-temperature damage to the components inside the satellite during use.

[0004] The above object of the application is achieved by the following technical scheme:

[0005] A sliding unfolding device for micro-nano satellites, comprising a satellite main body, one opposite side of one end of the satellite main body is hingedly connected with one solar wing plate, an elastic driving piece is arranged at the connection between the solar wing plate and the satellite main body, and the elastic driving piece can make the solar wing plate connected thereto in an unfolded state at the side edge of the satellite main body in a normal state.

[0006] A retractable support unit is mounted on the inner side of the bottom of the shell of the satellite main body, a horizontal linkage rod is connected to the top of the support unit, both ends of the horizontal linkage rod penetrate through the side walls of the satellite main body corresponding to the two solar wing plates, and both ends of the horizontal linkage rod are slidably connected to the shell of the satellite main body in the vertical direction.

[0007] The horizontal linkage rod is provided with a clamping joint at each end, and the two solar wings are provided with a clamping groove at the side close to the satellite body.

[0008] The upper side of the middle position of the horizontal linkage rod is provided with a remotely controlled adjusting unit, which can control the position of the horizontal linkage rod on the top of the supporting unit in the vertical direction.

[0009] Further, an iron base is fixedly installed at the middle position of the horizontal linkage rod, the adjusting unit is an electromagnet, and the electromagnet is fixedly installed in the shell of the satellite body and electrically connected with the control module and the battery module in the satellite body.

[0010] Further, the supporting unit comprises an outer sleeve, the lower end of the outer sleeve is fixedly connected with the bottom of the shell of the satellite body, an inner sleeve is coaxially arranged in the outer sleeve and they are in sliding connection, and the upper end of the inner sleeve is fixedly connected with the lower end of the iron base.

[0011] Further, a return spring is arranged in the inner sleeve, the lower end of the return spring is fixedly connected with the bottom of the shell of the satellite body, and the upper end of the return spring is fixedly connected with the inner side of the upper end of the inner sleeve.

[0012] Further, the clamping groove comprises two mutually symmetrical slide rails, each of the slide rails is provided with a socket at the side close to the other slide rail along the length direction of the slide rail, the side of the slide rail close to the hinged part of the satellite body and the solar wing is provided with a disengaging port, and an end plate is fixedly installed on the side opposite to the disengaging port.

[0013] Further, the clamping joint comprises two mutually symmetrical mounting plates, one end of each of the mounting plates is fixedly connected with one end of the horizontal linkage rod through a rivet, the other end of each of the mounting plates is fixedly connected with a limiting plate at the side away from each other, the limiting plate and the mounting plate connected therewith form a 90° angle, and the shape of the end away from each other of the limiting plate matches the shape in the slide rail.

[0014] Further, the length of the slide rail is greater than the length of the limiting plate.

[0015] Further, the elastic driving member is a torsion spring, and the torsion spring is installed on the hinge shaft of the hinged part of the satellite body and the solar wing.

[0016] In summary, the present application has at least one of the following beneficial technical effects:

[0017] Before the micro-nano satellite is put into the designated orbit, the solar wing plates can be clamped on the clamping joints at the two ends of the horizontal linkage rod through the clamping grooves, so that the micro-nano satellite as a whole does not occupy too much space in the carrying device. When the micro-nano satellite is put into the designated orbit, the ground control technician can drive the adjusting unit in the satellite main body shell to change the overall length of the supporting unit, so that the horizontal linkage rod with the two clamping joints simultaneously releases the restriction on the solar wing plates. The solar wing plates are reset quickly under the action of the elastic driving element until the two solar wing plates are in a fully unfolded state on both sides of the satellite main body. The unfolding device of the application not only has a simple structure and is convenient to control, but also does not generate excessive heat that may damage the internal components of the satellite during use, thereby effectively improving the safety and stability of the micro-nano satellite during unfolding. BRIEF DESCRIPTION OF DRAWINGS

[0018] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the following will briefly introduce the drawings needed to be used in the embodiments or the prior art description. Obviously, the drawings in the following description are some embodiments of the present application, and other drawings can also be obtained by those skilled in the art without creative labor.

[0019] Figure 1 is a structural schematic view of the solar wing plates on both sides of the satellite main body in a folded state;

[0020] Figure 2 is a structural schematic view of the solar wing plates on both sides of the satellite main body in a folded state; Figure 1 is a structural schematic view of the solar wing plates on both sides of the satellite main body in a folded state;

[0021] Figure 3 is an enlarged structural schematic view of A in Figure 2

[0022] Figure 4 is an internal structural schematic view of the satellite main body with the lower half of the shell and the supporting unit both cut open when the solar wing plates on both sides of the satellite main body are in a folded state;

[0023] Figure 5 is a structural schematic view of the solar wing plates on both sides of the satellite main body in an unfolded state;

[0024] Figure 6 is a structural schematic view of the solar wing plates on both sides of the satellite main body in an unfolded state, with the lower half of the shell of the satellite main body cut open.

[0025] ​1, satellite body; 2, solar wing plate; 3, support unit; 31, outer sleeve; 32, inner sleeve; 4, horizontal linkage rod; 5, clamping head; 51, mounting plate; 52, limiting plate; 6, clamping groove; 61, sliding rail; 62, socket; 63, disengagement port; 64, end plate; 7, iron base; 8, electromagnet; 9, return spring; 10, torsional spring; 11, fixed block; 12, fixed rod; 13, positioning plate. DETAILED DESCRIPTION

[0026] To make the purposes, technical solutions and advantages of the embodiments of the present application clearer, the technical solutions in the embodiments of the present application are described clearly and completely below. Obviously, the described embodiments are some of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative work fall within the scope of protection of the present application.

[0027] As shown in Figure 1 , Figure 4 , Figure 5 and Figure 6 , a sliding deployment device for a micro-nano satellite disclosed by the present application comprises a satellite body 1, one opposite side of one end of the satellite body 1 is hingedly connected with one solar wing plate 2, and an elastic driving member is arranged at the connection between the solar wing plate 2 and the satellite body 1. The elastic driving member can make the solar wing plate 2 connected therewith in an unfolded state at the side of the satellite body 1 in a normal state.

[0028] A retractable support unit 3 is mounted on the inner side of the bottom of the shell of the satellite body 1, and a horizontal linkage rod 4 is connected to the top of the support unit 3. The two ends of the horizontal linkage rod 4 respectively penetrate the side walls of the satellite body 1 corresponding to the two solar wing plates 2, and the two ends of the horizontal linkage rod 4 are slidably connected to the shell of the satellite body 1 in the vertical direction.

[0029] One clamping head 5 is arranged at each end of the horizontal linkage rod 4, and one clamping groove 6 is arranged on each side of the solar wing plate 2 close to the satellite body 1. The solar wing plate 2 is detachably connected to the clamping head 5 on the corresponding side through the clamping groove 6.

[0030] A remote control adjusting unit is arranged on the upper side of the horizontal linkage rod 4 at the middle position. The adjusting unit can control the position of the horizontal linkage rod 4 at the top of the support unit 3 in the vertical direction.

[0031] In the above embodiment, one of the opposite sides of the satellite main body 1 is provided with a solar wing plate 2, and the two solar wing plates 2 are hinged on the shell of the satellite main body 1, so that the solar wing plates 2 can be folded or unfolded on the satellite main body 1 as needed. When the micro-nano satellite is transported to the predetermined orbit by the carrier, the space in the carrier is often limited, and the solar wing plates 2 of the micro-nano satellite need to be folded on both sides of the satellite main body 1 during the process. At this time, the two solar wing plates 2 can be clamped on the clamping heads 5 at both ends of the horizontal linkage rod 4 through the clamping grooves 6 arranged on them, and the folded state of the solar wing plates 2 is maintained through the clamping heads 5 at both ends of the horizontal linkage rod 4, so as to ensure that the micro-nano satellite does not occupy too much space in the air carrier.

[0032] When the air carrier puts the micro-nano satellite into the predetermined orbit, the ground control technician can control the adjusting unit in the shell of the satellite main body 1 through remote wireless technology, and change the overall length of the supporting unit 3 by using the adjusting unit, such as making the overall length of the supporting unit 3 longer. The upper end of the supporting unit 3 will make the horizontal linkage rod 4 with the two clamping heads 5 simultaneously move upward in the vertical direction to escape from the clamping grooves 6 on the two solar wing plates 2, thereby releasing the restriction on the solar wing plates 2 (a through hole is arranged on the shell of the satellite main body 1 corresponding to the moving direction of the horizontal linkage rod 4, and the length of the through hole meets the needs of the moving range of the horizontal linkage rod 4 during use, so that the through hole arranged on the shell of the satellite main body 1 can ensure smooth movement during the movement of the horizontal linkage rod 4), and the solar wing plates 2 will be quickly reset under the action of the reset force of the elastic driving part until the two solar wing plates 2 are in a fully unfolded state on both sides of the satellite main body 1. Compared with the prior art, the unfolding device of the application not only has a simple structure and is convenient to operate, but also does not generate excessive heat that may damage the internal components of the satellite during use, thereby effectively improving the safety and stability of the micro-nano satellite during unfolding.

[0033] Further, as shown in Figure 4 and Figure 6 , an iron base 7 is fixedly installed at the middle position of the horizontal linkage rod 4, the adjusting unit is an electromagnet 8, and the electromagnet 8 is fixedly installed in the shell of the satellite main body 1. The electromagnet 8 is electrically connected with the control module and the battery module in the satellite main body 1.

[0034] In the above embodiment, when the solar wing plates 2 on both sides of the satellite main body 1 need to be unfolded, the ground operator can remotely control the electromagnet 8 to be powered on. After the electromagnet 8 is powered on, it can generate an attractive force on the iron base 7, so that the iron base 7 moves synchronously with the horizontal linkage rod 4 to the direction of the electromagnet 8. When the iron base 7 is adsorbed to the electromagnet 8, the clamping joints 5 at both ends of the horizontal linkage rod 4 are just separated from the clamping grooves 6 on the two solar wing plates 2. At this time, the elastic reset member at the hinge of the solar wing plate 2 and the satellite main body 1 will drive the solar wing plate 2 to unfold quickly. The main purpose of powering on the electromagnet 8 is to attract the horizontal linkage rod 4 with the two clamping joints 5 to release the restriction on the solar wing plate 2. Therefore, after the solar wing plate 2 is unfolded, the power supply to the electromagnet 8 can be stopped, which can effectively reduce the waste of the satellite main body 1 in the use process.

[0035] Further, as shown in Figure 4 and Figure 6 , the support unit 3 includes an outer sleeve 31, the lower end of the outer sleeve 31 is fixedly connected with the shell bottom of the satellite main body 1, the inner sleeve 32 is coaxially arranged in the inner sleeve 32 and is in sliding connection therebetween, and the upper end of the inner sleeve 32 is fixedly connected with the lower end of the iron base 7.

[0036] In the above embodiment, when the electromagnet 8 attracts the iron base 7 in the middle of the horizontal linkage rod 4, the iron base 7 will move together with the inner sleeve 32 to the direction of the electromagnet 8. The inner sleeve 32 is in sliding connection with the outer sleeve 31, and the lower end of the outer sleeve 31 is fixedly installed on the shell bottom of the satellite main body 1. Thus, during the movement of the iron base 7 to the direction of the electromagnet 8, the outer sleeve 31 can provide a guiding effect on the movement of the iron base 7 and the horizontal linkage rod 4 through the inner sleeve 32, so that the movement is stable enough. Even after the electromagnet 8 is powered off, the horizontal linkage rod 4 will not shake along its length direction.

[0037] Further, as shown in Figure 4 , the inner sleeve 32 is provided with a reset spring 9, the lower end of the reset spring 9 is fixedly connected with the shell bottom of the satellite main body 1, and the upper end of the reset spring 9 is fixedly connected with the inner side of the upper end of the inner sleeve 32.

[0038] In the above embodiment, after the electromagnet 8 completes the operation of pulling out the clamping joint 5 at both ends of the horizontal linkage rod 4 from the clamping groove 6 on the solar wing plate 2, in order to avoid the continuous waste of the micro-nano satellite's electric energy by the electromagnet 8, the electromagnet 8 can be powered off at this time. After the electromagnet 8 is powered off, the attractive force of the electromagnet 8 to the iron base 7 is lost, at this time the elastic force of the reset spring 9 reoccupies the dominant position, and the iron base 7 with the horizontal linkage rod 4 is driven to reset by the elastic force of the reset spring 9. The horizontal linkage rod 4 after resetting can limit its position through the reset spring 9 in the process of micro-nano satellite running, so that the risk of horizontal linkage rod 4 shaking in the vertical direction (the vertical direction in this application mainly refers to the height direction of the shell in the satellite main body 1) can be further reduced on the basis of the previous embodiment.

[0039] Further, as shown in Figure 2 and Figure 3 , the clamping groove 6 includes two mutually symmetrical slide rails 61, and each of the mutually close sides of the two slide rails 61 is provided with a socket 62 along the length direction thereof. The side of the slide rail 61 close to the hinged part of the solar wing plate 2 and the satellite main body 1 is provided with a disengagement port 63, and the side opposite to the disengagement port 63 of the slide rail 61 is fixedly installed with an end plate 64.

[0040] In the above embodiment, the two slide rails 61 are mutually symmetrical along the width direction of the solar wing plate 2 and have a spacing therebetween. The open side of the slide rail 61 close to the hinged part of the solar wing plate 2 and the satellite main body 1 is the disengagement port 63 thereof, which facilitates the clamping joint 5 at the end of the horizontal linkage rod 4 to exit the slide rail 61. The sockets 62 provided on the mutually close sides of the two slide rails 61 facilitate the clamping joint 5 at the end of the horizontal linkage rod 4 to slide along the length direction of the slide rail 61.

[0041] Further, as shown in Figure 2 and Figure 3 , the clamping joint 5 includes two mutually symmetrical mounting plates 51. One end of each of the two mounting plates 51 is fixedly connected with one end of the horizontal linkage rod 4 through a rivet. The mutually far apart sides of the other ends of the two mounting plates 51 are respectively fixedly connected with one limiting plate 52 each. The limiting plate 52 and the mounting plate 51 connected therewith form a 90° angle. The shapes of the mutually far apart ends of the two limiting plates 52 match the shapes in the slide rail 61.

[0042] In the above embodiment, the two limiting plates 52 can be stably connected between the mounting plate 51 and the horizontal linkage rod 4, so that the horizontal linkage rod 4 moves synchronously with all the limiting plates 52. The shape of the mutually distal sides of the limiting plates 52 matches the shape of the cross section of the slide rail 61, so that when the solar wing panel 2 of the micro-nano satellite is in the folded state, the mutually distal ends of the two limiting plates 52 on each side of the satellite main body 1 are inserted into the slide rail 61 at the corresponding position, thereby limiting and maintaining the folded state. When the horizontal linkage rod 4 moves in the vertical direction, the limiting plates 52 can stably move along the length direction of the slide rail 61 together with the horizontal linkage rod 4.

[0043] Further, as shown in Figure 4 , the length of the slide rail 61 is greater than the length of the limiting plate 52.

[0044] In the above embodiment, the micro-nano satellite inevitably shakes due to speed changes during transportation, which may cause the limiting plate 52 to accidentally fall out of the slide rail 61. In the present application, the length of the slide rail 61 is greater than the length of the limiting plate 52, so that even if the limiting plate 52 accidentally shakes in the vertical direction, it will not easily fall out of the slide rail 61, thereby ensuring that the micro-nano satellite maintains the folded state of the solar wing panel 2 during transportation.

[0045] Further, as shown in Figure 1 and Figure 5 , the elastic driving member is a torsion spring 10, which is installed on the hinge shaft of the satellite main body 1 and the solar wing panel 2.

[0046] In the above embodiment, the torsion spring 10 is a common spiral spring, as shown in Figure 5 , one end of the torsion spring 10 installed on the hinge shaft of the satellite main body 1 and the solar wing panel 2 is clamped below the fixed block 11 on the shell of the satellite main body 1, and the other end of the torsion spring 10 is clamped below the fixed rod 12 below the solar wing panel 2. When the solar wing panel 2 is folded on the side of the satellite main body 1, the two ends of the torsion spring 10 will rotate relative to each other, at which time the torsion spring 10 will generate a torque and a rotating force. After the clamping joints 5 at the two ends of the horizontal linkage rod 4 are removed from the corresponding slide rails 61 on the respective sides, the rotating force of the torsion spring 10 can be released, thereby driving the solar wing panel 2 to unfold on the side of the satellite main body 1. As shown in Figure 4 and Figure 6 , in actual use, in order to prevent the solar wing panel 2 from unfolding too much, a positioning plate 13 can be added outside the shell of the satellite main body 1, which can be abutted by the solar wing panel 2 when it is unfolded to the horizontal state, thereby ensuring that the unfolded solar wing panel 2 is in a stable horizontal state.

[0047] The implementation principle of the embodiment is as follows: during transportation of the micro-nano satellite, the two solar wing plates 2 are in a folded state at both sides of the satellite, at this time, the clamping joints 5 at both ends of the horizontal linkage rod 4 are respectively inserted into the slide rails 61 of the two solar wing plates 2, the inner sleeve 32 supporting the horizontal linkage rod 4 is also retracted into the outer sleeve 31 by the reset spring 9. After the micro-nano satellite is transported and released to the designated orbit, the ground operator can power on the electromagnet 8, after the electromagnet 8 is powered on, the attractive force generated by the electromagnet 8 to the ferrous base 7 is greater than the traction force of the reset spring 9 to the inner sleeve 32, the horizontal linkage rod 4 and the ferrous base 7, at this time, the ferrous base 7 will quickly move towards the electromagnet 8 with the horizontal linkage rod 4 and the inner sleeve 32, after the ferrous base 7 and the electromagnet 8 are attracted tightly, the clamping joints 5 at both ends of the horizontal linkage rod 4 are just detached from the clamping grooves 6 on the two solar wing plates 2, the torsional spring 10 installed on the hinge shaft of the solar wing plate 2 and the satellite main body 1 will quickly release the elastic force generated by folding, driving the two solar wing plates 2 to automatically adjust to the unfolded state at both sides of the satellite main body 1. After the solar wing plates 2 are unfolded, the electromagnet 8 can be powered off, the electromagnet 8 after being powered off no longer attracts the ferrous base 7, at this time, the reset spring 9 will reset the ferrous base 7 and the horizontal linkage rod 4 through the inner sleeve 32 by using the elastic force generated by deformation. Compared with the prior art, the unfolding device of the application does not generate excessive heat that may damage the internal components of the satellite during use, which can effectively improve the safety and stability of the micro-nano satellite during unfolding.

[0048] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the application, and not to limit them; although the application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that they can modify the technical solutions described in the foregoing embodiments, or make equivalent replacement for part or all of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the scope of the technical solutions of the embodiments of the application.

Claims

1. A sliding deployment device for micro-nano satellites, comprising a satellite body (1), characterized in that: The opposite side of one end of the satellite body (1) is respectively hinged with a solar wing plate (2), the connecting part of the solar wing plate (2) and the satellite body (1) is provided with an elastic driving element, the elastic driving element can make the solar wing plate (2) connected therewith in an unfolded state at the side of the satellite body (1) in normal state; The inside of the bottom of the shell of the satellite body (1) is provided with a telescopic supporting unit (3), the top of the supporting unit (3) is connected with a horizontal linkage rod (4), the two ends of the horizontal linkage rod (4) respectively penetrate the side walls of the satellite body (1) corresponding to the two solar wing plates (2), and the two ends of the horizontal linkage rod (4) are slidably connected with the shell of the satellite body (1) in the vertical direction; The two ends of the horizontal linkage rod (4) are respectively provided with a clamping joint (5), the side close to the satellite body (1) of the two solar wing plates (2) is respectively provided with a clamping groove (6), and the solar wing plate (2) is detachably connected with the clamping joint (5) on the corresponding side through the clamping groove (6). The upper side of the middle position of the horizontal linkage rod (4) is provided with a remotely controlled adjusting unit, and the adjusting unit can control the position of the horizontal linkage rod (4) on the top of the supporting unit (3) in the vertical direction.

2. The sliding deployment device for micro-nano satellites according to claim 1, characterized in that: The middle position of the horizontal linkage rod (4) is fixedly provided with an iron base (7), the adjusting unit is an electromagnet (8), the electromagnet (8) is fixedly installed in the shell of the satellite body (1), and the electromagnet (8) is electrically connected with a control module and a battery module in the satellite body (1).

3. The sliding deployment device for micro-nano satellites according to claim 2, characterized in that: The supporting unit (3) comprises an outer sleeve (31), the lower end of the outer sleeve (31) is fixedly connected with the bottom of the shell of the satellite body (1), an inner sleeve (32) is coaxially arranged in the inner sleeve (31) and is slidably connected with the outer sleeve (31), and the upper end of the inner sleeve (32) is fixedly connected with the lower end of the iron base (7).

4. The sliding deployment device for micro-nano satellites according to claim 3, characterized in that: The inner sleeve (32) is provided with a reset spring (9), the lower end of the reset spring (9) is fixedly connected with the bottom of the shell of the satellite body (1), and the upper end of the reset spring (9) is fixedly connected with the inner side of the upper end of the inner sleeve (32). 5.The sliding deployment device for micro-nano satellites according to claim 1, characterized in that: The clamping groove (6) comprises two mutually symmetrical slide rails (61), one socket (62) is arranged on the mutually close sides of the two slide rails (61) along the length direction of the slide rails (61), a disengaging port (63) is arranged on the side close to the hinged part of the solar wing plate (2) and the satellite body (1) of the slide rail (61), and an end plate (64) is fixedly installed on the side opposite to the disengaging port (63) of the slide rail (61).

6. The sliding deployment device for micro-nano satellites according to claim 5, characterized in that: The clamping joint (5) comprises two mutually symmetrical mounting plates (51), one end of the two mounting plates (51) is fixedly connected with one end of the horizontal linkage rod (4) through a rivet, and the other end of the two mounting plates (51) is fixedly connected with a limiting plate (52) on the side away from each other, and the limiting plate (52) and the mounting plate (51) connected therewith are at a 90° angle; the shape of the mutually away end of the two limiting plates (52) matches the shape in the slide rail (61).

7. The sliding deployment device for micro-nano satellites according to claim 6, characterized in that: The length of the slide rail (61) is greater than the length of the limiting plate (52).

8. The sliding deployment device for micro-nano satellites according to any one of claims 1-6, characterized in that: The elastic driving part is a torsion spring (10), and the torsion spring (10) is installed on the hinge shaft of the solar wing plate (2) and the satellite main body (1).

Images