Design method for on-orbit launch separation mechanism of miniature spacecraft

Abstract

The invention discloses a design method of an on-orbit launch separation mechanism of a miniature spacecraft. The separation mechanism based on the design method mainly comprises a main frame structure, a pull pin unlocking device, a memory alloy pull pin device, a tail charging port support and a protective skin. The main frame comprises a guide rail, a flange plate, a reinforcing ring and a tail baffle. The memory alloy pin puller and the pin pulling unlocking device are used for achieving locking of the miniature spacecraft in the separation mechanism before launching, the influence of vibration on the miniature spacecraft can be effectively reduced through cooperation with a tail baffle, and the miniature spacecraft can be prevented from colliding with a tail charging port of the separation mechanism. The separation mechanism designed through the design method is light in weight and high in space utilization rate, the separation reliability of the miniature spacecraft is high, and the separation guiding precision of the miniature spacecraft can be effectively improved.

Description

Technical Field

[0001] The invention relates to the technical field of satellite launch, in particular to a design method for an on-orbit launch and separation mechanism of a micro spacecraft suitable for rapid launch of a solid thruster. Background Art

[0002] At present, with the rapid development of aerospace technology, the initial separation speed that traditional micro-spacecraft separation mechanisms can provide is low, which cannot meet the mission requirements of some micro-spacecraft that require a higher initial separation speed. The use of solid thrusters for launch has the advantages of short response time and fast launch speed, and has gradually become a popular solution for the in-orbit separation of micro-spacecraft.

[0003] However, due to the characteristics of short action time and large specific impulse of solid thrusters, the micro-spacecraft will be accelerated to the set speed in a very short time. It is necessary to design a separation mechanism for this situation to constrain the movement of the micro-spacecraft during the ignition of the solid thrusters, so as to improve the launch and separation posture of the micro-spacecraft and improve the launch accuracy. Summary of the invention

[0004] Aiming at the on-orbit launch of a micro-spacecraft using a solid thruster, the present invention proposes a design method for an on-orbit launch and separation mechanism of a micro-spacecraft.

[0005] The technical solution to realize the present invention is: a design method for an on-orbit launch and separation mechanism of a micro-spacecraft, the steps are as follows:

[0006] Step 1: Determine the envelope and configuration of the separation mechanism according to the outer envelope of the micro-spacecraft, determine the gap and number of the guide rails, and the contact form with the guide button of the micro-spacecraft, and divide the guide rails from front to back into second-order guide rails and first-order guide rails.

[0007] Step 2: According to the measured thrust data of the solid thruster, the sliding length of the first-stage guide rail is determined by simulation. The sliding length should ensure that the rear guide button of the micro-spacecraft does not slide off the first-stage guide rail when the solid thruster ignition is completed.

[0008] Step 3: Based on the thrust deviation generated by the solid thruster, the height difference between the first-stage guide rail and the second-stage guide rail is determined by simulation. The height of the first-stage guide rail is greater than that of the second-stage guide rail. The height difference should ensure that the micro-spacecraft will not collide again after sliding off the first-stage guide rail and before separating from the separation mechanism.

[0009] Step 4: According to the overall structural strength and rigidity requirements of the separation mechanism, taking into account the mechanical interface between the separation mechanism and the satellite platform, and at the same time, in order to achieve no influence on the sliding effect of the guide rails and the controllable gap assembly of the guide rails, several reinforcement rings are arranged at intervals outside the four guide rails to form the frame body, from front to back, they are the front reinforcement ring, the first middle reinforcement ring, the second middle reinforcement ring, the third middle reinforcement ring, and the rear reinforcement ring. The cross-section and installation position of the reinforcement ring are determined under the premise of weight reduction, and the frame body is fixed to the satellite platform through the flange and the third middle reinforcement ring.

[0010] Step 5: Select and determine the locking and unlocking mechanisms of the separation mechanism to lock the micro-spacecraft before launch and unlock it without affecting the launch of the micro-spacecraft.

[0011] Step 6: To prevent the jet plume generated by the ignition of the solid thruster from entering the interior of the satellite platform and damaging components and equipment, a protective skin is designed based on the frame body, and the skin thickness is determined based on the structural strength requirements and overall weight reduction needs.

[0012] Step 7: To prevent cold welding, all moving parts in the separation mechanism are coated.

[0013] Compared with the prior art, the present invention has the following significant advantages:

[0014] (1) The designed separation mechanism guide rail adopts a two-stage simultaneous sliding mode. At the same time, the contact surface of the guide rail is coated with molybdenum disulfide. The front and rear guide keys on the micro-spacecraft slide off the corresponding guide rail sections at the same time, meeting the requirements of low-disturbance rapid launch of the micro-spacecraft and effectively improving the guiding effect of the separation of the micro-spacecraft.

[0015] (2) The design uses a memory alloy pin puller unlocking method instead of an initiator unlocking method, which effectively reduces the impact on the separation platform at the unlocking moment and reduces pollution. It replaces the traditional non-initiator electromagnetic unlocking method and effectively solves the problem of excessive residual magnetism in the traditional non-initiator solution.

[0016] (3) When the separation mechanism is not unlocked, the micro-spacecraft is held in place by the main pin inside the separation mechanism and is limited by the tail baffle, thereby effectively reducing the impact of vibration on the micro-spacecraft and effectively preventing the micro-spacecraft from excessively contacting the tail charging bracket and the charging contacts and causing a collision.

[0017] The above-described objects, features and advantages as well as the working principle of the present invention will be described and explained in detail below with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] Figure 1 The present invention is a simplified flowchart of the design method of the in-orbit launch and separation mechanism of a micro spacecraft.

[0019] Figure 2 It is a schematic diagram of the principle of a two-stage simultaneous sliding separation mechanism.

[0020] Figure 3 It is an overall schematic diagram of the micro spacecraft separation mechanism based on this design method.

[0021] Figure 4 It is a schematic diagram of the guide rail of the micro spacecraft separation mechanism based on this design method.

[0022] Figure 5 This is a schematic diagram of the memory alloy pin puller and its components selected based on this design method.

[0023] In the figure, 1. front reinforcement ring, 2. first middle reinforcement ring, 3. second middle reinforcement ring, 4. third middle reinforcement ring, 5. rear reinforcement ring, 6. flange, 7. pin puller, 8. memory alloy pin puller, 9. skin, 10. tail baffle, 11. guide rail, 12. charging bracket. DETAILED DESCRIPTION

[0024] The design method is described in detail below with reference to the accompanying drawings and specific implementation methods.

[0025] Combination Figure 1 to Figure 5 , a design method for an on-orbit launch and separation mechanism of a micro spacecraft, the steps are as follows:

[0026] Step 1: Determine the envelope and configuration of the separation mechanism according to the outer envelope of the micro-spacecraft, and determine the gap and quantity of the guide rails 11, as well as the contact form with the guide buttons of the micro-spacecraft.

[0027] Furthermore, since the solid thruster provides a large acceleration to the micro-spacecraft in a very short time when it is working (the ignition time of the micro-spacecraft is 20ms during launch, and the maximum acceleration is about 800m / s 2 ), in order to improve the separation accuracy of the separation mechanism, the guide rail configuration is designed to slide away in two stages at the same time, such as Figure 2 , Figure 4 As shown; according to the strength, stiffness and weight reduction requirements of the separation mechanism, the number of guide rails 11 is determined to be four, and each guide rail 11 is divided into a second-order guide rail and a first-order guide rail from front to back; the gap of the guide rail 11 is determined according to the gap matching of the existing spacecraft separation mechanism; in order to limit the rotation of the micro-spacecraft, right-angle guide buttons are used to cooperate with each other.

[0028] Step 2: According to the measured thrust data of the solid thruster, the sliding length of the first-stage guide rail is determined through ADAMS software simulation. This sliding length should ensure that the rear guide button of the micro-spacecraft does not slide off the first-stage guide rail when the solid thruster ignition is completed;

[0029] Furthermore, due to the uneven combustion of the solid thruster during ignition, a thrust eccentricity of about 20′ will be generated. At this stage, the separation mechanism should limit the rotation and circumferential movement of the micro-spacecraft.

[0030] A micro-spacecraft-separation mechanism model is established in ADAMS, and its actual mass and moment of inertia are given. The measured solid thruster thrust curve is applied to obtain the micro-spacecraft sliding distance at the end of solid thruster ignition. To achieve the above purpose, the first-stage guide rail length of the separation mechanism should be greater than the micro-spacecraft sliding length to ensure that the micro-spacecraft completes the solid thruster ignition task within the first-stage guide rail.

[0031] Step 3: Based on the thrust deviation generated by the solid thruster, determine the height difference between the first-stage guide rail and the second-stage guide rail through ADAMS software simulation. The height of the first-stage guide rail is greater than the height of the second-stage guide rail. The height difference should ensure that the micro-spacecraft will not collide again after sliding off the first-stage guide rail and before separating from the separation mechanism.

[0032] Furthermore, after determining the height difference, the lengths of the first-stage guide rail and the second-stage guide rail are continued, and the thrust eccentricity of the solid thruster is increased. Through ADAMS software simulation, the height difference between the first-stage guide rail and the second-stage guide rail is determined to ensure that the rear directional button of the micro spacecraft does not collide with the second-stage guide rail after separating from the first-stage guide rail, so as not to affect the separation attitude and launch accuracy of the micro spacecraft.

[0033] Step 4: According to the overall structural strength and rigidity requirements of the separation mechanism, taking into account the mechanical interface between the separation mechanism and the satellite platform, and at the same time, in order to achieve no influence on the sliding effect of the guide rail and the controllable gap assembly of the guide rail 11, a number of reinforcement rings are arranged at intervals outside the four guide rails 11 to form the frame body, from front to back, they are, front reinforcement ring 1, first middle reinforcement ring 2, second middle reinforcement ring 3, third middle reinforcement ring 4, and rear reinforcement ring 5. The cross-section and installation position of the reinforcement ring are determined under the premise of weight reduction, and the frame body is fixedly connected to the satellite platform through flange 6 and the third middle reinforcement ring 4.

[0034] Furthermore, the connection surface between the reinforcing ring and the guide rail is designed on both sides of the guide rail 11 to avoid affecting the flatness of the guide rail 11. At the same time, the connection surface between the reinforcing ring and the guide rail can achieve adjustable gap between the guide rail 11 and the micro spacecraft by adding ultra-thin gaskets.

[0035] Step 5: Select and determine the locking and unlocking mechanism design of the separation mechanism to lock the micro-spacecraft before launch and unlock it without affecting the launch of the micro-spacecraft.

[0036] Furthermore, the pin puller 7 and the memory alloy pin puller 8 in the patent CN114852380B are selected, and the specific structure is as follows: Figure 5As shown, the locking, unlocking and launching of the micro spacecraft are realized.

[0037] Furthermore, since the micro-spacecraft and the separation mechanism are clearance-fitted, and considering that the micro-spacecraft may collide with the guide rail and the main pin of the pin puller during transportation and launch, in order to alleviate the shear force on the main pin, a number of tail baffles 10 are installed at equal intervals at the tail of the frame body for limiting. This can effectively reduce the impact of vibration on the micro-spacecraft, and at the same time effectively prevent the charging contacts of the micro-spacecraft from excessively contacting and colliding with the charging bracket 12 at the tail of the frame body.

[0038] Step 6: To prevent the jet plume generated by the ignition of the solid thruster from entering the interior of the satellite platform and damaging components and equipment, a protective skin 9 is designed based on the frame body, and the thickness of the skin is determined based on the structural strength requirements and the overall weight reduction needs.

[0039] Step 7: To prevent cold welding, all movable parts in the separation mechanism (including the pin unlocker 7 and the guide rail 11) are subjected to coating treatment.

Claims

1. A design method for an on-orbit launch and separation mechanism of a micro-spacecraft, characterized in that: Here are the steps: Step 1: Determine the envelope and configuration of the separation mechanism according to the outer envelope of the micro-spacecraft, determine the gap and quantity of the guide rail (11), and the contact form with the guide button of the micro-spacecraft, and divide the guide rail (11) from front to back into a second-order guide rail and a first-order guide rail; Step 2: According to the measured thrust data of the solid thruster, the sliding length of the first-stage guide rail is determined by simulation. The sliding length should ensure that the rear guide button of the micro-spacecraft does not slide off the first-stage guide rail when the solid thruster ignition is completed; Step 3: According to the thrust deviation generated by the solid thruster, the height difference between the first-stage guide rail and the second-stage guide rail is determined by simulation. The height of the first-stage guide rail is greater than that of the second-stage guide rail. The height difference should ensure that the micro-spacecraft will not collide again after sliding off the first-stage guide rail and before separating from the separation mechanism. Step 4: According to the requirements of the overall structural strength and rigidity of the separation mechanism, taking into account the mechanical interface between the separation mechanism and the satellite-carrying platform, and in order to achieve the goal of not affecting the sliding effect of the guide rail and the controllable assembly of the clearance of the guide rail (11), a plurality of reinforcing rings are arranged at intervals outside the four guide rails (11) to form a frame body, which are, from front to back, a front reinforcing ring (1), a first middle reinforcing ring (2), a second middle reinforcing ring (3), a third middle reinforcing ring (4), and a rear reinforcing ring (5). The cross section and installation position of the reinforcing ring are determined under the premise of weight reduction, and the frame body is fixedly connected to the satellite-carrying platform through a flange (6) and a third middle reinforcing ring (4); Step 5: Select and determine the locking and unlocking mechanism of the separation mechanism to lock the micro-spacecraft before launch and unlock it without affecting the launch of the micro-spacecraft; Step 6: To prevent the jet plume generated by the ignition of the solid thruster from entering the interior of the satellite platform and damaging the components and equipment, a protective skin (9) is designed based on the frame body, and the thickness of the skin is determined based on the structural strength requirements and the overall weight reduction requirements; Step 7: To prevent cold welding, all moving parts in the separation mechanism are coated.

2. The design method of the micro-spacecraft on-orbit launch and separation mechanism according to claim 1 is characterized in that: In step one, in order to improve the separation accuracy of the separation mechanism, the guide rail configuration is designed to slide away simultaneously in two stages; according to the strength, rigidity and weight reduction requirements of the separation mechanism, the number of guide rails (11) is determined to be four, and each guide rail (11) is divided into a second-order guide rail and a first-order guide rail from front to back; the clearance of the guide rail (11) is determined according to the clearance fit of the existing spacecraft separation mechanism; and in order to limit the rotation of the micro-spacecraft, right-angle guide buttons are used to cooperate with each other.

3. The design method of the micro-spacecraft on-orbit launch and separation mechanism according to claim 2 is characterized in that: In step 2, according to the measured thrust data of the solid thruster, the sliding length of the first-stage guide rail is determined by simulation. This sliding length should ensure that the rear guide button of the micro-spacecraft does not slide off the first-stage guide rail when the solid thruster ignition is completed. The details are as follows: Since the solid thruster burns unevenly during ignition, thrust eccentricity will occur. At this stage, the separation mechanism should limit the rotation and circumferential movement of the micro-spacecraft. A micro-spacecraft-separation mechanism model is established in the simulation software, and its actual mass and moment of inertia are given. The measured thrust curve of the solid thruster is applied to obtain the micro-spacecraft sliding distance at the end of the solid thruster ignition.

4. The design method of the micro-spacecraft on-orbit launch and separation mechanism according to claim 3 is characterized in that: In step 2, the length of the first-stage guide rail of the separation mechanism should be greater than the sliding length of the micro-spacecraft to ensure that the micro-spacecraft completes the solid thruster ignition task within the first-stage guide rail.

5. The design method of the micro-spacecraft on-orbit launch and separation mechanism according to claim 4 is characterized in that: In step 3, the height difference between the first-stage guide rail and the second-stage guide rail is determined. The height of the first-stage guide rail is greater than that of the second-stage guide rail. The height difference should ensure that the micro-spacecraft does not collide again after sliding off the first-stage guide rail and before separating from the separation mechanism. The details are as follows: After determining the height difference, continue to calculate the length of the first-stage guide rail and the second-stage guide rail, increase the thrust eccentricity of the solid thruster, and determine the height difference between the first-stage guide rail and the second-stage guide rail through simulation to ensure that the rear directional button of the micro spacecraft does not collide with the second-stage guide rail after leaving the first-stage guide rail, so as not to affect the separation attitude and launch accuracy of the micro spacecraft.

6. The design method of the micro-spacecraft on-orbit launch and separation mechanism according to claim 5 is characterized in that: In step 4, a plurality of reinforcing rings are arranged at intervals outside the four guide rails (11) to form a frame body, specifically as follows: the connecting surfaces between the reinforcing rings and the guide rails are designed on both sides of the guide rails (11) to avoid affecting the flatness of the guide rails (11), and at the same time, the connecting surfaces between the reinforcing rings and the guide rails are adjusted by adding ultra-thin gaskets to achieve an adjustable gap between the guide rails (11) and the micro-spacecraft.

7. The design method of the micro-spacecraft on-orbit launch and separation mechanism according to claim 6 is characterized in that: In step 5, since the micro-spacecraft and the separation mechanism are clearance-matched, considering that the micro-spacecraft may collide with the guide rail and the main pin of the pin puller during transportation and launch, in order to alleviate the shear force on the main pin, a plurality of tail baffles (10) are installed at equal intervals at the rear of the frame body for limiting, which effectively reduces the impact of vibration on the micro-spacecraft and effectively prevents the charging contacts of the micro-spacecraft from excessively contacting and colliding with the charging bracket (12) at the rear of the frame body.

8. The design method of the micro-spacecraft on-orbit launch and separation mechanism according to claim 7 is characterized in that: In step seven, the movable parts include a pin unlocker (7) and a guide rail (11).

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