Design method of on-orbit separation mechanism of micro spacecraft

Abstract

The application discloses a design method of a micro-spacecraft on-orbit launch separation mechanism, and a separation mechanism based on the design method mainly comprises a main frame structure, a pin pulling unlocker, a memory alloy pin pulling 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 pulling device and the pin pulling unlocker are used to lock the micro-spacecraft in the separation mechanism before launch, and the tail baffle can effectively reduce the influence of vibration on the micro-spacecraft and prevent the micro-spacecraft from colliding with the tail charging port of the separation mechanism. The separation mechanism designed by the design method has the advantages of light weight, high space utilization, high micro-spacecraft separation reliability and high micro-spacecraft separation guiding precision.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of satellite launching technology, in particular to a design method of an on-orbit launching separation mechanism for a micro-satellite launched by a solid thruster. BACKGROUND

[0002] At present, with the rapid development of space technology, the initial separation speed provided by the traditional micro-satellite separation mechanism is small, which cannot meet the task requirements of some micro-satellites that need higher initial separation speed. The solid thruster has the advantages of short response time and fast launching speed, and gradually becomes a popular scheme for on-orbit separation of micro-satellites.

[0003] The solid thruster will lift the micro-satellite to a set speed in a very short time due to its short action time and large specific impulse, and the separation mechanism needs to be designed for this situation to constrain the movement of the micro-satellite during the ignition of the solid thruster, so as to improve the launching and separating posture of the micro-satellite and improve the launching precision. SUMMARY

[0004] The present application provides a design method of an on-orbit launching separation mechanism for a micro-satellite launched by a solid thruster.

[0005] The technical solution of the present application is as follows:

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

[0007] Step two: simulate to determine the sliding length of the first-order guide rail according to the measured thrust data of the solid thruster, which should ensure that the rear guide knob of the micro-satellite does not slide out of the first-order guide rail when the solid thruster is ignited.

[0008] Step three: simulate to determine the height difference between the first-order guide rail and the second-order guide rail according to the thrust deviation generated by the solid thruster, the height of the first-order guide rail is greater than the height of the second-order guide rail, and the height difference should ensure that the micro-satellite does not collide again after sliding out of the first-order guide rail and before separating from the separation mechanism.

[0009] Step four: according to the overall structure strength and rigidity requirement of the separation mechanism, the mechanical interface of the separation mechanism and the satellite platform is considered, and the clearance assembly of the guide rail is controlled without affecting the sliding effect of the guide rail, a plurality of reinforcing rings are arranged outside the four guide rails to form a frame main body, and the reinforcing rings are sequentially arranged from front to back, front reinforcing ring, first middle reinforcing ring, second middle reinforcing ring, third middle reinforcing ring and rear reinforcing ring, the cross section and mounting position of the reinforcing ring are determined under the premise of weight reduction, and the frame main body is fixedly connected with the satellite platform through the flange plate and the third middle reinforcing ring.

[0010] Step five: the locking and unlocking mechanism of the separation mechanism is selected and determined, so that the micro-satellite is locked before launch and is not affected after unlocking.

[0011] Step six: in order to prevent the jet plume generated by the ignition of the solid thruster from entering the inside of the satellite platform and damaging the equipment, a protective skin is designed according to the frame main body, and the thickness of the skin is determined according to the structural strength requirement and the overall weight reduction requirement.

[0012] Step seven: in order to prevent cold welding, all movable parts in the separation mechanism are coated.

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

[0014] (1) The designed separation mechanism guide rail is a two-section type and slides away 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-satellite slide away at the same time and correspond to the guide rail sections, the demand of low disturbance and rapid launch of the micro-satellite is met, and the guiding effect of the micro-satellite separation is effectively improved.

[0015] (2) The design adopts a memory alloy puller unlocking method instead of a pyrotechnic unlocking method, which effectively reduces the impact on the separation platform at the unlocking moment and reduces pollution. Instead of the traditional non-pyrotechnic electromagnetic unlocking method, the problem of excessive residual magnetism of the traditional non-pyrotechnic scheme is effectively solved.

[0016] (3) When the separation mechanism is not unlocked, the micro-satellite is stopped by the main pin in the separation mechanism, and the micro-satellite is limited by the tail baffle, which effectively reduces the influence of vibration on the micro-satellite, and effectively prevents the micro-satellite from colliding with the tail charging support and charging contact due to excessive contact.

[0017] The purposes, features and advantages described above and the working principle of the present application will be described and explained in detail with reference to the accompanying drawings. BRIEF DESCRIPTION OF DRAWINGS

[0018] Figure 1 It is a design method flowchart of a micro-satellite on-orbit launch separation mechanism.

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

[0020] Figure 3 is the overall schematic diagram of micro-satellite separation mechanism based on the design method.

[0021] Figure 4 is the guide rail schematic diagram of micro-satellite separation mechanism based on the design method.

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

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

[0024] The design method will be described in detail below in combination with the drawings and specific embodiments.

[0025] In combination Figures 1-5 , a design method of micro-satellite on-orbit launch separation mechanism, the steps are as follows:

[0026] Step one: according to the outer envelope of micro-satellite, determine the envelope and configuration of separation mechanism, determine the gap and number of guide rail 11, and the contact form with micro-satellite guide knob.

[0027] Further, since the solid thruster will provide a large acceleration to the micro-satellite in a very short time when working (the micro-satellite is ignited for 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 as two-stage simultaneous sliding, as shown in Figure 2 、 Figure 4 According to the strength, stiffness and weight reduction requirements of the separation mechanism, it is determined that the number of guide rails 11 is four, each guide rail 11 is divided into second-order guide rail and first-order guide rail from front to back; according to the existing spacecraft separation mechanism gap matching, the gap of guide rail 11 is determined; in order to limit the rotation of micro-satellite, the right-angle guide knob is used.

[0028] Step two: according to the measured thrust data of solid thruster, determine the sliding length of first-order guide rail through ADAMS software simulation, which should ensure that the rear guide knob of micro-satellite does not slide off the first-order guide rail when the solid thruster is ignited.

[0029] Further, due to the uneven combustion of the solid thruster when ignited, a thrust eccentricity of about 20' is generated, and in this stage the separation mechanism should limit the rotation and circumferential movement of the micro-satellite.

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

[0031] Step three: According to the thrust deviation generated by the solid thruster, the height difference between the first guide rail and the second guide rail is determined through ADAMS software simulation, the height of the first guide rail is greater than the height of the second guide rail, and the height difference should ensure that the micro-satellite does not collide again after slipping out of the first guide rail and before separating from the separation mechanism.

[0032] Further, after determining the height difference, the length of the first guide rail and the second guide rail is continued, the solid thruster thrust eccentricity is increased, and the height difference between the first guide rail and the second guide rail is determined through ADAMS software simulation, to ensure that the micro-satellite rear orientation knob does not collide with the second guide rail after leaving the first guide rail, so as to avoid affecting the separation attitude of the micro-satellite and the launch precision.

[0033] Step four: According to the overall structural strength and stiffness requirements of the separation mechanism, the mechanical interface of the separation mechanism and the satellite platform is considered, and at the same time, the gap assembly of the guide rail 11 is controlled without affecting the guide rail slip effect, a number of reinforcing rings are arranged outside the four guide rails 11 to form a frame main body, from front to back, they are front reinforcing ring 1, first middle reinforcing ring 2, second middle reinforcing ring 3, third middle reinforcing ring 4, and 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 main body is fixed to the satellite platform through the flange plate 6 and the third middle reinforcing ring 4.

[0034] Further, the reinforcing ring and the guide rail connection surface are designed on both sides of the guide rail 11 to avoid affecting the flatness of the guide rail 11, and the reinforcing ring and the guide rail connection surface can realize the gap adjustment between the guide rail 11 and the micro-satellite by increasing the ultra-thin gasket.

[0035] Step five: The locking and unlocking mechanism design of the separation mechanism is selected and determined to lock the micro-satellite before launch and unlock without affecting the launch of the micro-satellite.

[0036] Further, the pin extractor 7 and the memory alloy pin extractor 8 in the patent CN114852380B are selected, and the specific structure is as follows Figure 5As shown, the micro spacecraft is locked and unlocked for launching.

[0037] Further, since the micro spacecraft and the separation mechanism are gap-fitted, the micro spacecraft may collide with the guide rail and the main pin of the pin extractor during transportation and launching. To relieve the shear force on the main pin, several tail baffles 10 are installed at intervals at the tail of the frame body to limit the position, which can effectively reduce the influence of vibration on the micro spacecraft, and effectively prevent the collision of the micro spacecraft charging contact with the charging support 12 at the tail of the frame body.

[0038] Step six: To prevent the jet plume generated by the ignition of the solid thruster from entering the inside of the satellite platform and damaging the components and devices, a protective skin 9 is designed according to the frame body, and the thickness of the skin is determined according to the requirements of structural strength and overall weight reduction.

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

Claims

1. A method for designing an on-orbit launch separation mechanism for a microspacecraft, characterized by, The steps are as follows: Step one: according to the outer envelope of the micro-satellite, the envelope and configuration of the separation mechanism are determined, the gap and number of the guide rails (11) are determined, and the contact form with the micro-satellite guide knob is determined, the guide rails (11) are divided into second-order guide rails and first-order guide rails from front to back; Step two: according to the measured thrust data of the solid thruster, the sliding length of the first-order guide rail is determined by simulation, which should ensure that the micro-satellite rear guide knob does not slide off the first-order guide rail at the end of the solid thruster ignition; Step three: according to the thrust deviation generated by the solid thruster, the height difference between the first-order guide rail and the second-order guide rail is determined by simulation, the height of the first-order guide rail is greater than that of the second-order guide rail, and the height difference should ensure that the micro-satellite does not collide after sliding off the first-order guide rail and before separating from the separation mechanism; Step four: according to the overall structural strength and stiffness requirements of the separation mechanism, the mechanical interface of the separation mechanism and the satellite platform is considered, and the sliding effect of the guide rail and the gap of the guide rail (11) are controlled, a number of reinforcing rings are arranged outside the four guide rails (11) to form a frame main body, which is sequentially arranged from front to back as 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 main body is fixed to the satellite platform through the flange (6), the third middle reinforcing ring (4); Step five: select and determine the locking and unlocking mechanism of the separation mechanism to lock the micro-satellite before launch and unlock it without affecting the launch of the micro-satellite; Step six: In order to prevent the jet plume generated by the solid thruster from entering the inside of the satellite platform and damaging the equipment, a protective skin (9) is designed according to the frame main body, and the skin thickness is determined according to the structural strength requirement and overall weight reduction requirement; Step seven: In order to prevent cold welding, all moving parts in the separation mechanism are coated.

2. The method of designing an on-orbit release mechanism for a microspacecraft according to claim 1, wherein In step one, in order to improve the separation accuracy of the separation mechanism, the guide rail configuration is designed as two sections simultaneously sliding off; according to the strength, stiffness and weight reduction requirements of the separation mechanism, the number of guide rails (11) is determined as four, each guide rail (11) is divided into second-order guide rails and first-order guide rails from front to back; according to the existing spacecraft separation mechanism gap fit, the gap of the guide rail (11) is determined; in order to limit the rotation of the micro-satellite, the straight guide knob is used.

3. The method of designing an on-orbit release mechanism for a microspacecraft according to claim 2, wherein In step two, according to the measured thrust data of the solid thruster, the sliding length of the first-order guide rail is determined by simulation, which should ensure that the micro-satellite rear guide knob does not slide off the first-order guide rail at the end of the solid thruster ignition, which is as follows: Due to the uneven combustion of the solid thruster during ignition, thrust eccentricity will occur, and the separation mechanism should limit the rotation and circumferential movement of the micro-satellite at this stage; In the simulation software, a micro-satellite-separation mechanism model is established, and its actual mass and moment of inertia are given, and the measured solid thruster thrust curve is applied to obtain the sliding distance of the micro-satellite at the end of the solid thruster ignition.

4. The method of claim 3, wherein the micro-satellite on-orbit launch separation mechanism is designed to be separated from the launch vehicle by a separation mechanism. In step two, the length of the first stage guide rail of the separation mechanism should be greater than the sliding length of the micro-satellite, ensuring that the micro-satellite completes the solid propeller ignition task in the first stage guide rail.

5. The method of designing an on-orbit release mechanism for a microspacecraft of claim 4, wherein: In step three, the height difference between the first stage guide rail and the second stage guide rail is determined, and 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-satellite does not collide again after sliding out of the first stage guide rail and before separating from the separation mechanism. The specific steps are as follows: After determining the height difference, continue to increase the length of the first stage guide rail and the second stage guide rail to increase the eccentricity of the solid propeller thrust. Through simulation, the height difference between the first stage guide rail and the second stage guide rail is determined to ensure that the micro-satellite does not collide with the second stage guide rail after the rear orientation knob is separated from the first stage guide rail, so as to avoid affecting the separation attitude of the micro-satellite and the launch precision.

6. The method of designing an on-orbit release mechanism for a microspacecraft according to claim 5, wherein In step four, a number of reinforcing rings are arranged outside the four guide rails (11) to form a frame body. The specific steps are as follows: the connecting surface of 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), and the connecting surface of the reinforcing ring and the guide rail is adjusted by increasing the ultra-thin gasket to realize the gap adjustment between the guide rail (11) and the micro-satellite.

7. The method of designing an on-orbit release mechanism for a microspacecraft according to claim 6, wherein In step five, since the micro-satellite and the separation mechanism are gap-fitted, the micro-satellite may collide with the guide rail and the main pin of the pin extractor during transportation and launch. To alleviate the shear force on the main pin, a number of tail baffles (10) are installed at the tail of the frame body at equal intervals to limit the position, effectively reducing the influence of vibration on the micro-satellite, and effectively preventing the micro-satellite charging contact from colliding with the charging support (12) at the tail of the frame body.

8. The method of designing an on-orbit release mechanism for a microspacecraft according to claim 7, wherein, In step seven, the movable parts include the pin extractor unlocker (7) and the guide rail (11).

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