A parallel spacecraft ejection and separation device

Abstract

The present invention discloses a parallel spacecraft ejection separation device. The ejection separation device includes an installation bottom plate and an upper installation plate arranged opposite to each other; two support beams are arranged in parallel and fixedly installed on the installation bottom plate; a multi-link mechanism is hinged between each support beam and the upper installation plate; the two multi-link mechanisms are symmetrically arranged, and a plurality of parallel coupling shafts are connected therebetween; a sliding member is slidably installed between the two support beams along the support beams and is hinged to one of the link rods in the multi-link mechanism; a parallel cylinder is fixedly installed on the installation bottom plate; the piston rod of the parallel cylinder is fixedly connected to the sliding member; a damper is fixedly installed on each support beam. The above ejection separation device can realize the adjustability of the ejection stroke, ejection load, and ejection separation angle, assist the pneumatic separation to increase the launch safety, and increase the universality of the ejection mechanism, and can eject and separate different second-stage spacecrafts.

Description

Technical Field

[0001] The present invention relates to the technical field of aerospace launch separation, and particularly to a parallel spacecraft ejection separation device. Background Art

[0002] Reusable spacecraft can significantly reduce the cost of launching payloads into space. Technically, reusable spacecraft can be basically divided into three categories: the first is the spacecraft category, with the basic technical characteristics of vertical launch and vertical recovery; the second is the space shuttle category, with the basic technical characteristics of vertical launch and horizontal recovery; the third is the aerospace plane category, with the basic technical characteristics of horizontal takeoff, horizontal landing and recovery.

[0003] Among them, the piggyback two-stage orbital aerospace vehicle is a research hotspot. The first stage is a carrier aircraft, and the second stage is an aerospace plane. The aerospace plane is carried on the back of the carrier aircraft for horizontal takeoff and landing and high-altitude launch. The configuration is similar to that of the Soviet An-225 carrying "Buran" or Boeing 747 carrying the US space shuttle, but the latter two are only used for the transfer transportation of the space shuttle, while the two-stage orbital aerospace vehicle is used for the launch of the aerospace plane.

[0004] Existing technologies for spacecraft separation include: explosive separation devices, which achieve spacecraft separation by detonating separation explosives or gunpowder. This method can provide a relatively high separation force, but due to the use of explosives, there are certain safety risks, which may cause accidents, damage to the spacecraft or even accidents, and it is not easy to control the separation process; mechanical separation devices, which use mechanical structures such as spiral springs and pins to achieve spacecraft separation. This method is relatively simple and reliable, but careful design of the separation mechanism is required in terms of separation force and control accuracy; hydraulic or pneumatic separation devices, which use hydraulic or pneumatic systems to apply pressure when separation is required to achieve spacecraft separation. This method can provide relatively high control accuracy and reliability, but requires corresponding hydraulic or pneumatic systems; electromagnetic separation devices, which achieve spacecraft separation by using electromagnetic force or electromagnetic actuators. This method has relatively high control accuracy and reliability, and does not require the use of explosives or hydraulic systems, but may require additional power supplies and control circuits.

[0005] There is relatively little research on the air launch separation device of the piggyback two-stage orbital aerospace vehicle and publicly available dedicated separation mechanisms. Therefore, it is necessary to design a simple and efficient launch separation device.

[0006] Although the prior art provides some solutions for spacecraft separation, for the separation of parallel spacecraft, at present, pneumatic separation is mainly adopted. Pneumatic separation has high requirements for aerodynamics. Relying solely on pneumatic separation has low reliability. It is necessary to assist with an active separation force on the basis of pneumatic separation design to increase launch safety and reliability. At this time, it is necessary to design a catapult mechanism that can adjust the ejection stroke, control the ejection load size, and adjust the separation angle. Summary of the Invention

[0007] The present invention provides a catapult separation device for parallel spacecraft, which can achieve adjustable ejection stroke, ejection load, and ejection separation angle, assist pneumatic separation to increase launch safety, and increase the universality of the catapult mechanism, and can catapult and separate different second-stage spacecraft.

[0008] The present invention adopts the following specific technical solutions:

[0009] A catapult separation device for parallel spacecraft, the catapult separation device includes a mounting base plate, two support beams, a parallel cylinder, a sliding member, two multi-link mechanisms, an upper mounting plate, and a damper;

[0010] The mounting base plate is disposed opposite to the upper mounting plate; the mounting base plate is used to connect the first-stage spacecraft; the upper mounting plate is used to support the second-stage spacecraft;

[0011] The two support beams are arranged in parallel and fixedly installed on the top surface of the mounting base plate;

[0012] A multi-link mechanism is hinged between each support beam and the upper mounting plate; the two multi-link mechanisms are symmetrically arranged and are connected by a plurality of parallel coupling shafts;

[0013] The sliding member is slidably installed between the two support beams along the length extension direction of the support beams and is hinged to a link in each multi-link mechanism;

[0014] The parallel cylinder is horizontally and fixedly installed on the top surface of the mounting base plate; the piston rod of the parallel cylinder is fixedly connected to the sliding member and is used to drive the sliding member to slide along the support beam, so as to drive the upper mounting plate to move in the vertical direction through the expansion and contraction of the multi-link mechanism;

[0015] A damper is fixedly installed on each support beam, and the damper is used to buffer the sliding member when the second-stage spacecraft is catapulted and separated.

[0016] Furthermore, the sliding member is provided with a locking mechanism for locking it to the support beam.

[0017] Further, it also includes two hydraulic support columns for supporting the upper mounting plate when the multi-link mechanism is in a contracted state and separating during ejection;

[0018] One of the hydraulic support columns is fixedly installed on each of the support beams;

[0019] The hydraulic support column is provided with a fixed connection structure at the top;

[0020] When the hydraulic support column supports the upper mounting plate, the fixed connection structure is used to fixedly connect with the upper mounting plate.

[0021] Further, the fixed connection structure is magnetic attraction.

[0022] Further, the parallel cylinders include a first cylinder and a second cylinder fixedly installed on the mounting base plate;

[0023] The first piston rod of the first cylinder and the second piston rod of the second cylinder move synchronously and are both fixedly connected to the sliding member.

[0024] Further, the two multi-link mechanisms are a first multi-link mechanism and a second multi-link mechanism that are mirror-symmetrical;

[0025] The first multi-link mechanism includes a first link, a second link, a third link, a fourth link, a fifth link, and a sixth link;

[0026] The second multi-link mechanism includes a seventh link opposite to the first link, an eighth link opposite to the second link, a ninth link opposite to the third link, a tenth link opposite to the fourth link, an eleventh link opposite to the fifth link, and a twelfth link opposite to the sixth link;

[0027] The coupling shafts include a first coupling shaft, a second coupling shaft, a third coupling shaft, and a fourth coupling shaft;

[0028] The first link, the second link, the third link, the fourth link, the fifth link, and the sixth link are hinged to form a double trapezoidal structure, where: the bottom ends of the first link and the third link are spaced apart and are both hinged to the support beam; the top end of the first link, the bottom end of the second link, and one end of the fifth link are rotationally connected by the first coupling shaft; the top end of the third link, the other end of the fifth link, and the middle of the fourth coupling shaft are rotationally connected by the second coupling shaft; the top end of the second link and one end of the sixth link are rotationally connected by the third coupling shaft and then hinged to the upper mounting plate; the other end of the sixth link and the top end of the fourth link are rotationally connected by the fourth coupling shaft and then hinged to the upper mounting plate; the bottom end of the fourth link is hinged to the sliding member.

[0029] Further, the lengths of the first link, the second link, the third link, and the fourth link satisfy the following formula:

[0030] L3 + L4 > L1 + L2 > L3 + L4 / 2;

[0031] Where, L1 is the length of the first link, L2 is the length of the second link, L3 is the length of the third link, and L4 is the length of the fourth link.

[0032] Further, the sliding member and the support beam are in concave-convex fit.

[0033] Further, the support beam is provided with a guiding groove extending along its length direction;

[0034] The sliding member is provided with a protrusion that is in shape fit with the guiding groove.

[0035] Further, both the support beam and the upper mounting plate are provided with trunnions for connecting the multi-link mechanism;

[0036] The parallel cylinder is a single-stage actuating cylinder.

[0037] Beneficial effects:

[0038] 1. The parallel spacecraft ejection and separation device of the present invention includes a mounting base plate for connecting the first-stage spacecraft, an upper mounting plate for supporting the second-stage spacecraft, two parallel support beams, a multi-link mechanism hinged between the support beams and the upper mounting plate, a sliding member slidably engaged with the support beams, a parallel cylinder fixedly mounted on the mounting base plate, and a damper fixedly mounted on the support beams; the piston rod of the parallel cylinder is fixedly connected to the sliding member and is used to drive the sliding member to slide along the support beams, so as to drive the upper mounting plate to move in the vertical direction through the expansion and contraction of the multi-link mechanism; the damper is used to buffer the sliding member; for the parallel spacecraft ejection and separation device with the above structure, by adjusting the charging pressure of the parallel cylinder, the moving distance of the piston rod, and the multi-link mechanism, the adjustment of the ejection force, ejection stroke, and ejection separation angle can be respectively realized. Therefore, the above parallel spacecraft ejection and separation device can achieve adjustable ejection stroke, ejection load, and ejection separation angle, assist pneumatic separation to increase launch safety, increase the universality of the ejection mechanism, and can eject and separate different second-stage spacecraft.

[0039] 2. The parallel spacecraft ejection and separation device of the present invention adopts a parallel single-stage actuating cylinder, which has the characteristics of simple structure and reliable recovery. Moreover, the parallel cylinder is horizontally placed, and when it contracts, the height, volume, and occupied space of the ejection and separation device are low.

[0040] 3. The parallel spacecraft ejection and separation device of the present invention converts the horizontal force of the parallel cylinder into a vertical force on the second-stage spacecraft through the sliding cooperation between the support beams and the sliding member and the expansion and contraction of the multi-link mechanism. It has a compact structure, a large ejection stroke, and good force transmission stiffness. At the same time, the overall horizontal stiffness of the ejection and separation device is improved by the fifth link and the eleventh link of the multi-link mechanism, reducing the horizontal deformation during the ejection process. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] Figure 1 is a three-dimensional structural schematic diagram of the parallel spacecraft ejection and separation device of the present invention;

[0042] Figure 2 is a structural schematic diagram of another perspective of the parallel spacecraft ejection and separation device of the present invention;

[0043] Figure 3 is a structural schematic diagram of another perspective of the parallel spacecraft ejection and separation device of the present invention;

[0044] Figure 4 is a structural schematic diagram when the multi-link mechanism is in a contracted state;

[0045] Figure 5 is a schematic diagram of the use state of the parallel spacecraft ejection and separation device of the present invention;

[0046] Figure 6For Figure 2 Partial enlarged structural schematic diagram of part A in

[0047] Figure 7 Assembly structural schematic diagram of the sliding component and the support beam;

[0048] Wherein, 1 - mounting base plate, 2 - support beam, 3 - parallel cylinders, 4 - sliding component, 5 - multi-link mechanism, 6 - upper mounting plate, 7 - hydraulic support column, 8 - damper, 9 - first-stage spacecraft, 10 - second-stage spacecraft, 31 - first cylinder, 32 - second cylinder, 311 - first piston rod, 321 - second piston rod, 511 - first link, 512 - second link, 513 - third link, 514 - fourth link, 515 - fifth link, 516 - sixth link, 521 - seventh link, 522 - eighth link, 523 - ninth link, 524 - tenth link, 525 - eleventh link, 526 - twelfth link, 531 - first coupling shaft, 532 - second coupling shaft, 533 - third coupling shaft, 534 - fourth coupling shaft Specific embodiments

[0049] Next, the technical solutions in the embodiments of the present invention will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0050] As Figure 1 、 Figure 2 and Figure 3 shown in the structure, the embodiment of the present invention provides a parallel spacecraft ejection separation device, and the ejection separation device includes a mounting base plate 1, two support beams 2, parallel cylinders 3, a sliding component 4, two multi-link mechanisms 5, an upper mounting plate 6 and a damper 8;

[0051] The mounting base plate 1 and the upper mounting plate 6 are oppositely arranged in the vertical direction; the upper mounting plate 6 is arranged at intervals on the top of the mounting base plate 1; both the mounting base plate 1 and the upper mounting plate 6 can be flat plates; as Figure 5 shown, the mounting base plate 1 is used to connect the first-stage spacecraft 9; the upper mounting plate 6 is used to support the second-stage spacecraft 10; the second-stage spacecraft 10 is connected to the ejection separation device through the upper mounting plate 6;

[0052] As Figure 1 and Figure 2As shown in the figure, two support beams 2 are arranged in parallel and fixedly installed on the top surface of the mounting base plate 1; a sliding space for the sliding member 4 is formed between the two support beams 2; the support beams 2 can be fixedly connected to the mounting base plate 1 by connection methods such as welding and bolts; the sliding member 4 and the support beams 2 are in concave-convex fit; as Figure 7 shown, the support beam 2 is provided with a guiding groove 21 extending along its length direction; the sliding member 4 is provided with a protrusion that matches the shape of the guiding groove 21; alternatively, a guiding protrusion can be provided on the support beam 2, and a sliding groove that matches the shape of the guiding protrusion can be provided on the sliding member 4; the two support beams 2 are of symmetric structure and are provided with guiding grooves for guiding the sliding member 4, and the upper surface of the support beam 2 is provided with trunnions for connecting the multi-link mechanism 5; the trunnions for connecting the multi-link mechanism 5 can be adjusted in position in the horizontal and vertical directions according to the ejection requirements to change the ejection effect;

[0053] A multi-link mechanism 5 is hinged between each support beam 2 and the upper mounting plate 6; the two multi-link mechanisms 5 are symmetrically arranged, and a plurality of parallel coupling shafts are connected between them; both the support beam 2 and the upper mounting plate 6 are provided with trunnions for connecting the multi-link mechanism 5, and both the upper mounting plate 6 and the support beam 2 can be connected to the multi-link mechanism 5 through the trunnions;

[0054] The sliding member 4 is slidably installed between the two support beams 2 along the length extension direction of the support beam 2 and is hinged to one link in each multi-link mechanism 5; the sliding member 4 can also be provided with a locking mechanism for locking it to the support beam 2. When the ejection separation device is in the retracted state, the sliding member 4 is locked to the support beam 2 by the locking mechanism. Locking the sliding member 4 to the support beam 2 through the locking mechanism can improve the stability of the multi-link mechanism 5, the upper mounting plate 6, and the second-stage spacecraft 10 installed on the upper mounting plate 6; when ejection is performed, the locking mechanism is unlocked;

[0055] As Figure 1 and Figure 2 shown, the parallel cylinder 3 is horizontally and fixedly installed on the top surface of the mounting base plate 1; the piston rod of the parallel cylinder 3 is fixedly connected to the sliding member 4 and is used to drive the sliding member 4 to slide along the support beam 2, so as to drive the upper mounting plate 6 to move in the vertical direction through the expansion and contraction of the multi-link mechanism 5; the parallel cylinder 3 includes a first cylinder 31 and a second cylinder 32 fixedly installed on the mounting base plate 1. Both the first cylinder 31 and the second cylinder 32 can be connected to the top surface of the mounting base plate 1 by bolts. The first piston rod 311 of the first cylinder 31 and the second piston rod 321 of the second cylinder 32 are both fixedly connected to the sliding member 44 and act synchronously, such as: by threaded connection, welding, riveting, etc., and remain parallel to meet the synchronism; the parallel cylinder 3 adopts a single-stage actuating cylinder;

[0056] A damper 8 is fixedly installed on each support beam 2. The damper 8 is used to buffer the sliding member 4 when the second-stage spacecraft 10 is catapulted and separated. The damper 8 can be a spring damper. For example, Figure 6 as shown, the damper 8 can be installed at the end of the guiding groove of the support beam 2 for buffering and braking the sliding member 4.

[0057] The above-mentioned parallel spacecraft catapult separation device includes a mounting base plate 1 for connecting the first-stage spacecraft 9, an upper mounting plate 6 for supporting the second-stage spacecraft 10, two parallel support beams 2, a multi-link mechanism 5 hinged between the support beam 2 and the upper mounting plate 6, a sliding member 4 slidably engaged with the support beam 2, a parallel cylinder 3 fixedly installed on the mounting base plate 1, and a damper 8 fixedly installed on the support beam 2; the piston rod of the parallel cylinder 3 is fixedly connected to the sliding member 4 and is used to drive the sliding member 4 to slide along the support beam 2, so as to drive the upper mounting plate 6 to move in the vertical direction through the expansion and contraction of the multi-link mechanism 5; the damper 8 is used to buffer the sliding member 4; for the parallel spacecraft catapult separation device with the above structure, the catapult force is adjusted by adjusting the charging air pressure of the parallel cylinder 3, the catapult stroke is adjusted by the movement distance of the piston rod, and the adjustable catapult angle is realized by adjusting the height of the mounting trunnion for hinging the first link 511 and the seventh link 521 on the support beam 2 in the vertical direction, that is, by adjusting the height of the mounting trunnion, the height of the upper mounting plate 6 on one side of the first link 511 is raised, while the height of the other side of the upper mounting plate 6 is lowered, and the catapult inclination angle is controlled by the height difference between the two sides, so as to realize the adjustment of the catapult angle. Therefore, the above-mentioned parallel spacecraft catapult separation device can realize the adjustable catapult stroke, catapult load, and catapult separation angle, assist the pneumatic separation to increase the launch safety, increase the universality of the catapult mechanism, and can catapult and separate different second-stage spacecraft 10.

[0058] At the same time, due to the adoption of the parallel single-stage actuating cylinder, the whole catapult separation device has the characteristics of simple structure and reliable recovery. Moreover, the parallel cylinder 3 is horizontally placed, and when it contracts, the height of the catapult separation device is low, the volume is small, and the occupied space is small.

[0059] In a specific embodiment, the above catapult separation device further includes two hydraulic support columns 7 for supporting the upper mounting plate 6 when the multi-link mechanism 5 is in a contracted state and separating during catapult; a hydraulic support column 7 is fixedly installed on each support beam 2; the hydraulic support column 7 is provided with a fixed connection structure at the top; when the hydraulic support column 7 supports the upper mounting plate 6, the fixed connection structure is used to be fixedly connected to the upper mounting plate 6. The bottom end of the hydraulic support column 7 is fixedly connected to the support beam 2 by bolts, welding, etc.; for example, Figure 4As shown, it plays a supporting role when the ejection mechanism is in a contracted state. When supporting, the upper supporting surface of the hydraulic support column 7 can be fixedly connected to the upper mounting plate 6 through magnetic attraction, a locking mechanism, etc., and separated during ejection.

[0060] By supporting the upper mounting plate 6 with two hydraulic support columns 7, the weight borne by the multi-link mechanism 5 can be reduced, which is beneficial to protecting the multi-link mechanism 5 while improving the supporting stability. At the same time, the fixed connection structure on the hydraulic support column 7 is used to realize the fixed connection with the upper mounting plate 6, further improving the stability and reliability of the second-stage spacecraft 10.

[0061] Furthermore, as Figure 1 and Figure 3 shown, the two multi-link mechanisms 5 are a first multi-link mechanism and a second multi-link mechanism that are mirror-symmetrical; among them:

[0062] The first multi-link mechanism includes a first link 511, a second link 512, a third link 513, a fourth link 514, a fifth link 515, and a sixth link 516;

[0063] The second multi-link mechanism includes a seventh link 521 opposite to the first link 511, an eighth link 522 opposite to the second link 512, a ninth link 523 opposite to the third link 513, a tenth link 524 opposite to the fourth link 514, an eleventh link 525 opposite to the fifth link 515, and a twelfth link 526 opposite to the sixth link 516;

[0064] The coupling shafts include a first coupling shaft 531, a second coupling shaft 532, a third coupling shaft 533, and a fourth coupling shaft 534; the synchronous operation of the first multi-link mechanism and the second multi-link mechanism is realized through the connection of the coupling shafts;

[0065] The first link 511, the second link 512, the third link 513, the fourth link 514, the fifth link 515, and the sixth link 516 are hinged to form a double-trapezoid structure, where: the bottom ends of the first link 511 and the third link 513 are spaced apart and both are hinged to the support beam 2; the top end of the first link 511, the bottom end of the second link 512, and one end of the fifth link 515 are rotationally connected through the first coupling shaft 531; the top end of the third link 513, the other end of the fifth link 515, and the middle of the fourth coupling shaft 534 are rotationally connected through the second coupling shaft 532; the top end of the second link 512 and one end of the sixth link 516 are rotationally connected through the third coupling shaft 533 and then hinged to the upper mounting plate 6; the other end of the sixth link 516 and the top end of the fourth link 514 are rotationally connected through the fourth coupling shaft 534 and then hinged to the upper mounting plate 6; the bottom end of the fourth link 514 is hinged to the sliding member 4.

[0066] During ejection, the parallel cylinders 3 drive the sliding member 4 to slide along the support beam 2 through the piston rod. The sliding member 4 drives the bottom ends of the fourth link 514 and the tenth link 524 to move synchronously. The fourth link 514 and the tenth link 524 drive the third link 513 and the ninth link 523 to rotate synchronously around the trunnion on the support beam 2 through the second coupling shaft 532. At the same time, the first link 511 and the seventh link 521 are driven to rotate around the trunnion on the support beam 2 through the fifth link 515 and the eleventh link 525. Through the coordinated movement of the multi-link mechanism 5, the upper mounting plate 6 is driven to move in the vertical direction.

[0067] The above-mentioned ejection separation device for parallel spacecrafts converts the horizontal force of the parallel cylinders 3 into a vertical force on the second-stage spacecraft 10 through the sliding fit between the support beam 2 and the sliding member 4 and the expansion and contraction of the multi-link mechanism 5. It has a compact structure, a large ejection stroke, and good force transmission stiffness. At the same time, the overall horizontal stiffness of the ejection separation device is improved through the fifth link 515 and the eleventh link 525 of the multi-link mechanism 5, and the deformation in the horizontal direction during the ejection process is reduced.

[0068] To adjust the ejection angle, the lengths of the first link 511, the second link 512, the third link 513, and the fourth link 514 satisfy the following formula:

[0069] L3 + L4 > L1 + L2 > L3 + L4 / 2;

[0070] Where, L1 is the length of the first link 511, L2 is the length of the second link 512, L3 is the length of the third link 513, and L4 is the length of the fourth link 514;

[0071] Similarly, since the first multi-link mechanism 5 and the second multi-link mechanism 5 are mirror-symmetrical, the lengths of the seventh link 521, the eighth link 522, the ninth link 523, and the tenth link 524 satisfy the following formula:

[0072] L9 + L 10 >L7 + L8>L9 + L 10 / 2;

[0073] Where, L7 is the length of the seventh link 521, L8 is the length of the eighth link 522, L9 is the length of the ninth link 523, and L 10 is the length of the tenth link 524;

[0074] Since the sum of the lengths of the first link 511 and the second link 512 is less than the sum of the lengths of the third link 513 and the fourth link 514 and greater than half of the sum of the lengths of the third link 513 and the fourth link 514, the upper mounting plate 6 also undergoes a slight rotational movement during the vertical movement, thereby providing a lifting angle and angular velocity to the second-stage spacecraft 10 to achieve the adjustment of the ejection angle.

[0075] The working process of the above-mentioned parallel spacecraft ejection separation device includes:

[0076] The ejection separation device is locked by the locking mechanism of the sliding member 4, and the entire ejection separation device is driven by a cylinder to contract and extend;

[0077] When locked, the entire ejection separation device is in a contracted state, as Figure 4 shown, with a small height, small volume and small occupied space;

[0078] At the beginning of ejection, the locking mechanism of the sliding member 4 is unlocked. After unlocking, the sliding member 4 moves horizontally along the guiding groove of the support beam 2 driven by the cylinder. The sliding member 4 drives the fourth link 514 and the tenth link 524 to move through the upper connecting trunnion. The fourth link 514 and the tenth link 524 drive the entire multi-link mechanism 5 to extend through the coordinated action of the mechanism, pushing the upper mounting plate 6 to move vertically upward and accompanied by a small rotation. The sliding member 4 relies on the damper 8 for buffer braking at the end stage of the ejection phase;

[0079] After the ejection is completed, the entire ejection separation device can be retracted under the drive of the cylinder, and the process is opposite to the aforementioned extension process.

[0080] Obviously, those skilled in the art can make various changes and modifications to the embodiments of the present invention without departing from the spirit and scope of the present invention. Thus, if these modifications and variations of the present invention fall within the scope of the claims of the present invention and its equivalent technologies, the present invention is also intended to include these changes and modifications.

Claims

1. A parallel spacecraft ejection and separation device, characterized in that It includes a mounting base plate, two support beams, a parallel cylinder, a sliding component, two multi-link mechanisms, an upper mounting plate, and a damper; The mounting base plate is disposed opposite to the upper mounting plate; the mounting base plate is used to connect the first-stage spacecraft; the upper mounting plate is used to support the second-stage spacecraft; The two support beams are arranged in parallel and fixedly mounted on the top surface of the mounting base plate; One of the multi-link mechanisms is hinged between each of the support beams and the upper mounting plate; the two multi-link mechanisms are symmetrically arranged and are connected by a plurality of parallel coupling shafts; The sliding component is slidably mounted between the two support beams along the length extension direction of the support beam and is hinged to one of the links in each multi-link mechanism; The parallel cylinder is horizontally and fixedly mounted on the top surface of the mounting base plate; the piston rod of the parallel cylinder is fixedly connected to the sliding component and is used to drive the sliding component to slide along the support beam, so as to drive the upper mounting plate to move in the vertical direction through the expansion and contraction of the multi-link mechanism; One damper is fixedly mounted on each of the support beams, and the damper is used to buffer the sliding component when the second-stage spacecraft ejects and separates; The two multi-link mechanisms are the first multi-link mechanism and the second multi-link mechanism that are mirror-symmetrical; The first multi-link mechanism includes a first link, a second link, a third link, a fourth link, a fifth link, and a sixth link; The second multi-link mechanism includes a seventh link that is opposite to the first link, an eighth link that is opposite to the second link, a ninth link that is opposite to the third link, a tenth link that is opposite to the fourth link, an eleventh link that is opposite to the fifth link, and a twelfth link that is opposite to the sixth link; The coupling shafts include a first coupling shaft, a second coupling shaft, a third coupling shaft, and a fourth coupling shaft; The first link, the second link, the third link, the fourth link, the fifth link, and the sixth link are hinged to form a double-trapezoid structure, where: the bottom ends of the first link and the third link are spaced apart and are both hinged to the support beam; the top end of the first link, the bottom end of the second link, and one end of the fifth link are rotationally connected through the first coupling shaft; the top end of the third link, the other end of the fifth link, and the middle of the fourth coupling shaft are rotationally connected through the second coupling shaft; the top end of the second link and one end of the sixth link are rotationally connected through the third coupling shaft and then hinged to the upper mounting plate; the other end of the sixth link and the top end of the fourth link are rotationally connected through the fourth coupling shaft and then hinged to the upper mounting plate; the bottom end of the fourth link is hinged to the sliding component; The lengths of the first link, the second link, the third link, and the fourth link satisfy the following formula: L3 + L4 > L1 + L2 > L3 + L4 / 2; Wherein, L1 is the length of the first link, L2 is the length of the second link, L3 is the length of the third link, and L4 is the length of the fourth link.

2. The parallel spacecraft ejection separation device according to claim 1, characterized in that, The sliding member is provided with a locking mechanism for locking it to the support beam.

3. The parallel spacecraft ejection separation device according to claim 1, characterized in that, Also included are two hydraulic support columns for supporting the upper mounting plate when the multi-link mechanism is in a retracted state and for separating during ejection; A hydraulic support column is fixedly installed on each of the support beams; The hydraulic support column is provided with a fixed connection structure at the top end; When the hydraulic support column supports the upper mounting plate, the fixed connection structure is used to be fixedly connected to the upper mounting plate.

4. The parallel spacecraft ejection separation device according to claim 3, wherein, The fixed connection structure is magnetic attraction.

5. The parallel spacecraft ejection separation device according to claim 1, characterized in that The parallel cylinders include a first cylinder and a second cylinder fixedly mounted on the mounting base plate; The first piston rod of the first cylinder and the second piston rod of the second cylinder move synchronously and are both fixedly connected to the sliding component.

6. The parallel spacecraft ejection separation device according to claim 1, characterized in that, The sliding component and the support beam are matched with each other in a concave-convex manner.

7. The parallel spacecraft ejection separation device according to claim 6, characterized in that The support beam is provided with a guide groove extending along its length direction; The sliding component is provided with a protrusion which matches the shape of the guide groove.

8. The parallel spacecraft ejection separation device according to any one of claims 1-7, characterized in that, The support beam and the upper mounting plate are both provided with trunnions for connecting the multi-link mechanism; The parallel cylinder is a single-stage actuating cylinder.

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