Design methods, storage media, and computer equipment for near-field safety of spacecraft separation tests

Abstract

This invention provides a design method for near-field safety in satellite-launcher separation tests. The method involves acquiring the geometric and kinematic parameters of the test specimens; establishing a geometric model of the satellite and its launch vehicle support; analyzing the differences in operating conditions under ideal and unexpected states based on the kinematic parameters; determining the energy separation state during the ground test based on experimental realism and safety; analyzing the positional relationship between the satellite and its support during separation under unexpected conditions using the geometric model; and adjusting the height of the lower edge of the support from the ground protective pad until the minimum distance exceeds the safe separation distance. Thus, this invention enables rapid assessment of near-field safety in satellite-launcher separation tests under complex boundary and unexpected conditions, saving manpower and resources and reducing experimental risks.

Description

Technical Field

[0001] This invention relates to the field of satellite and rocket development technology, and in particular to a design method, storage medium, and computer equipment for near-field safety during satellite-rocket separation tests. Background Technology

[0002] The satellite-rocket locking and separation device is a key structure used for connecting, unlocking, and separating the satellite from the rocket. A common strap-type satellite-rocket locking and separation device consists of two straps connected in a ring by two explosive bolts. A V-shaped locking block is installed on the device, which locks the upper frame of the satellite support and the lower frame of the satellite. According to the flight sequence requirements, the control system issues a separation signal, the explosive bolts detonate, and the V-shaped locking block rapidly detaches from the docking frame under the elastic strain energy accumulated in the straps and docking frame. The satellite then detaches from the rocket under the action of the separation spring, thus achieving satellite-rocket separation.

[0003] The satellite-launcher ground docking and separation test is a crucial joint operation test project in the spacecraft development process. The differences between ground tests and flight tests introduce safety risks during separation. In some unexpected situations, such as unreleased explosive bolts or V-shaped locking blocks failing to disengage, collisions and damage have occurred between the launch vehicle support and spacecraft equipment during separation. However, current ground-based satellite-launcher separation tests rely solely on early engineering experience, and related dynamic simulations have failed to effectively analyze complex and unexpected conditions. This results in potential separation test hazards remaining into later stages of the mission, potentially leading to damage to the satellite and launch vehicle or even mission failure.

[0004] Therefore, conducting near-field safety design for satellite-rocket separation under extreme operating conditions is a necessary part of the satellite-rocket joint design work during the design phase. However, existing technologies do not provide near-field safety assessment methods for satellite-rocket separation ground tests. Summary of the Invention

[0005] The purpose of this invention is to provide a design method, storage medium, and computer equipment for near-field safety in space-rocket separation experiments, which can quantitatively analyze the impact of separation energy on separation safety under the differences between space and ground environments.

[0006] To achieve the above objectives, the present invention provides a design method for near-field safety in satellite-rocket separation experiments, comprising the following steps:

[0007] Obtain the geometric parameters of the test specimens for the satellite-rocket separation test; wherein, the geometric parameters include at least the radius of the satellite-rocket docking frame, the equipment envelope size of the satellite equipment extending into the satellite support, the extension height of the separation spring of the launch satellite support, the height of the satellite support, the radius of the lower edge of the satellite support, the height of the lower edge of the satellite support from the ground protective pad, and the safe distance for satellite-rocket separation;

[0008] Obtain the kinematic parameters corresponding to the test specimens of the satellite-rocket separation test; wherein, the kinematic parameters include satellite mass, satellite moment of inertia, launch satellite support mass, launch satellite support moment of inertia, launch vehicle final stage mass, launch vehicle final stage moment of inertia, and initial separation resultant force of the separation spring;

[0009] A geometric model of the satellite and its launch vehicle support is established based on the aforementioned geometric parameters;

[0010] Based on the geometric and kinematic parameters, the relative separation velocities of the satellite-rocket separation test specimen under ideal conditions are analyzed and calculated: the first relative separation velocity of the satellite-rocket in flight, the second relative separation velocity of the satellite-rocket in the energy-enabled state during ground testing, and the third relative separation velocity of the satellite-rocket in the energy-disabled state during ground testing. Based on the comparison between the first relative separation velocity, the second relative separation velocity, and the third relative separation velocity, the energy separation state during ground testing that meets the requirements for experimental authenticity is determined.

[0011] Based on the geometric and kinematic parameters, the relative rotational angular velocity of the satellite-rocket separation test specimen under unexpected conditions is analyzed and calculated as follows: the first satellite-rocket relative rotational angular velocity in flight state, the second satellite-rocket relative rotational angular velocity in ground test process with energy separation enabled, and the third satellite-rocket relative rotational angular velocity in ground test process with energy separation disabled. By comparing the second satellite-rocket relative rotational angular velocity with the third satellite-rocket relative rotational angular velocity, the ground test process energy separation state that meets the test safety requirements is determined.

[0012] Based on the geometric model, the positional relationship between the satellite and the satellite support during the separation process under unexpected operating conditions is analyzed, and the minimum distance between the satellite and the satellite support is obtained based on the positional relationship.

[0013] If the minimum spacing is less than the safe spacing between the satellite and the launch vehicle, the height of the lower edge of the satellite support from the ground protective pad is adjusted, and the minimum spacing is recalculated until the minimum spacing is greater than the safe spacing between the satellite and the launch vehicle.

[0014] Furthermore, the relative separation velocity of the first satellite and rocket is calculated based on the following formula:

[0015]

[0016] The relative separation velocity of the second satellite and rocket is calculated based on the following formula:

[0017]

[0018] The relative separation velocity of the third satellite-rocket is calculated based on the following formula:

[0019]

[0020] Wherein, m1 is the mass of the satellite, m2 is the mass of the launch satellite support, m3 is the mass of the final stage of the launch vehicle, F0 is the initial separation force of the separation spring, h is the extension height of the separation spring of the launch satellite support, and g is the acceleration due to gravity.

[0021] Furthermore, the relative rotational angular velocity of the first star-rocket is calculated based on the following formula:

[0022]

[0023] The relative rotational angular velocity of the second satellite-rocket is calculated based on the following formula:

[0024]

[0025] The relative rotational angular velocity of the third satellite-rocket is calculated based on the following formula:

[0026]

[0027] Wherein, I1 is the moment of inertia of the satellite, I2 is the moment of inertia of the launch satellite support, I3 is the moment of inertia of the final stage of the launch vehicle, F0 is the initial separation force of the separation spring, m2 is the mass of the launch satellite support, h is the extension height of the separation spring of the launch satellite support, and g is the acceleration due to gravity.

[0028] Furthermore, the step of analyzing the positional relationship between the satellite and the satellite support during the separation process under unexpected operating conditions based on the geometric model, and obtaining the minimum distance between the satellite and the satellite support based on the positional relationship, includes:

[0029] The process of separating the satellite from its support under unexpected operating conditions is defined in two stages: the first stage involves the satellite support rotating rigidly around the side connected to the satellite as a first fixed end under gravity until the lower edge of the satellite support contacts the ground protective pad; the second stage involves the satellite support rotating rigidly around the point where the lower edge of the satellite support contacts the ground protective pad as a second fixed end, the first fixed end being released from constraint, and the satellite support rotating rigidly around the second fixed end under gravity until the satellite support is in complete contact with the ground protective pad.

[0030] Based on the positional relationship between the satellite and the satellite support in the first and second stages of the geometric model, the minimum distance between the satellite and the satellite support is extracted.

[0031] Additionally, a storage medium and a computer device are provided, the storage medium being used to store a computer program for executing the above-described design method for near-field safety of the spacecraft separation test.

[0032] The computer device includes a storage medium, a processor, and a computer program stored on the storage medium and executable on the processor. When the processor executes the computer program, it implements the above-described design method for near-field safety of the star-rocket separation test.

[0033] The design method for near-field safety of spacecraft separation tests described in this invention can achieve rapid assessment of near-field safety of spacecraft separation tests under complex boundary and unexpected operating conditions based on motion geometry interference analysis and rigid body dynamics equations, saving manpower and resources and reducing test risks. Attached Figure Description

[0034] Figure 1 A flowchart illustrating the steps of a design method for near-field safety in a satellite-rocket separation test according to an embodiment of the present invention;

[0035] Figure 2 A detailed flowchart of the design method for near-field safety of the star-rocket separation test provided in an embodiment of the present invention;

[0036] Figure 3 This is a schematic diagram of the geometric model of the design method for near-field safety of the star-rocket separation test provided in an embodiment of the present invention.

[0037] Figure 4 for Figure 3 The diagram shown illustrates the separation motion of the geometric model before the height adjustment of the ground protective pad along the lower edge of the satellite support.

[0038] Figure 5 for Figure 3 The diagram shown illustrates the separation motion of the geometric model after adjusting the height of the satellite support along the ground protective pad. Detailed Implementation

[0039] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the invention.

[0040] It should be noted that references to "an embodiment," "embodiment," "example embodiment," etc., in this specification refer to the described embodiment including specific features, structures, or characteristics, but not every embodiment must include these specific features, structures, or characteristics. Furthermore, such expressions do not refer to the same embodiment. Moreover, when describing specific features, structures, or characteristics in conjunction with embodiments, whether or not explicitly described, it is indicated that incorporating such features, structures, or characteristics into other embodiments is within the knowledge of those skilled in the art.

[0041] Furthermore, certain terms are used in the specification and subsequent claims to refer to specific components or parts. Those skilled in the art will understand that manufacturers may use different names or terms to refer to the same component or part. This specification and subsequent claims do not distinguish components or parts by differences in name, but rather by differences in function. The terms "comprising" and "including" used throughout the specification and subsequent claims are open-ended and should be interpreted as "including but not limited to".

[0042] To address the shortcomings of current star-rocket separation tests that rely solely on engineering experience and whose dynamic simulations fail to effectively analyze complex and unexpected conditions, this invention provides a design method for near-field safety in star-rocket separation tests. This method, based on motion geometry interference analysis and rigid body motion equations, quantitatively analyzes the impact of separation energy on separation safety under the differences between the Earth and space environments, and provides a near-field safety design method for unexpected conditions.

[0043] Before describing the embodiments of the present invention, the various working conditions and test requirements of the satellite-rocket test targeted by the present invention are explained as follows:

[0044] The ideal working condition for satellite-rocket separation is defined as follows: no unexpected conditions occur during the satellite-rocket separation process, and the relative separation angular velocity of the satellite and rocket is zero.

[0045] The unexpected condition of satellite-rocket separation is defined as: an unexpected condition in which the tape jams during the satellite-rocket separation process, and there is a relative separation angular velocity between the satellite and the rocket.

[0046] Ground test authenticity requirements: Under ideal working conditions, provided that the separation energy mechanical interface is fully verified, the relative separation velocity of the satellite and rocket in the ground test should be as close as possible to the flight state;

[0047] Ground test safety requirements: Under unexpected operating conditions, the relative rotational angular velocity of the satellite and rocket during ground tests should be as small as possible.

[0048] Figure 1 This invention illustrates a design method for near-field safety in a spacecraft-rocket separation test, comprising the following steps:

[0049] S101: Obtain the geometric parameters of the test specimen for the satellite-rocket separation test; wherein, the geometric parameters include at least the radius of the satellite-rocket docking frame, the equipment envelope size of the satellite equipment extending into the satellite support, the extension height of the separation spring of the launch satellite support, the height of the satellite support, the radius of the lower edge of the satellite support, the height of the lower edge of the satellite support from the ground protective pad, and the safe distance for satellite-rocket separation. The test specimen for the satellite-rocket separation test is the test piece, including equipment such as a satellite, a launch satellite support, and a satellite-rocket locking and separation device; in step S101, the geometric parameters of the test specimen for the satellite-rocket separation test can be determined in the early design stage. In this embodiment, the geometric parameters specifically include the radius R0 of the satellite-rocket docking frame, the equipment envelope size of the satellite equipment extending into the satellite support (such as the extension length L of the orbital control engine, the lower edge radius r of the orbital control engine nozzle, etc.), the extension height h of the separation spring of the launch satellite support, the height H of the satellite support, the radius R1 of the lower edge of the satellite support, the height D of the lower edge of the satellite support from the ground protective pad, and the safe distance for satellite-rocket separation S.

[0050] The following is a specific application example, in which the specific values ​​of the geometric parameters of the test specimen for the star-rocket separation test are shown in Table 1;

[0051] Table 1:

[0052] Serial Number Geometric parameters Value / mm 1 <![CDATA[Radius R0 of the satellite-rocket docking frame]]> 595 2 Rail control engine insertion length L 440 3 The lower edge radius r of the track-controlled engine nozzle 150 4 The satellite launch support separation spring extends to a height h. 150 5 Satellite support height H 650 6 <![CDATA[The radius R1 of the lower edge of the satellite bracket]]> 875 7 The lower edge of the satellite support is at the same height (D) as the ground protective pad. 600 8 Star-rocket separation safety distance S 200

[0053] S102: Obtain the kinematic parameters corresponding to the test specimens for the satellite-rocket separation test; wherein, the kinematic parameters include satellite mass, satellite moment of inertia, launch satellite support mass, launch satellite support moment of inertia, launch vehicle final stage mass, launch vehicle final stage moment of inertia, and initial separation resultant force of the separation spring; the satellite moment of inertia and the launch vehicle support moment of inertia are both moments of inertia about one side of the docking frame; in step S102, the kinematic parameters corresponding to the test specimens for the satellite-rocket separation test can be determined in the early design stage.

[0054] Based on the above specific application examples, the specific values ​​of the kinematic parameters corresponding to the test specimens in the star-rocket separation test are shown in Table 2;

[0055] Table 2:

[0056] Serial Number kinematic parameters numerical values 1 <![CDATA[Satellite mass m1]]> 4300kg 2 <![CDATA[Satellite moment of inertia I1]]> <![CDATA[30000kg.m 2 ]]> 3 <![CDATA[Carrier satellite bracket mass m2]]> 45kg 4 <![CDATA[Carrier satellite bracket moment of inertia I2]]> <![CDATA[25kg.m 2 ]]> 5 <![CDATA[Final stage mass m3]]> 4200kg 6 <![CDATA[Inertia I3 of the upper stage of the launch vehicle]]> <![CDATA[67000kg.m 2 ]]> 7 <![CDATA[Initial separation resultant force F0 of the separation spring]]> 5880N

[0057] S103: Establish a geometric model of the satellite and the launch vehicle support based on the aforementioned geometric parameters; the geometric model of the satellite and the launch vehicle support established based on the above geometric parameters in this embodiment is as follows: Figure 3 As shown.

[0058] S104: Based on the geometric and kinematic parameters, analyze and calculate the relative separation velocity of the first satellite and rocket in flight, the relative separation velocity of the second satellite and rocket in the energy-enabled state during ground testing, and the relative separation velocity of the third satellite and rocket in the energy-disabled state during ground testing, under ideal conditions. Based on the comparison between the relative separation velocities of the first, second, and third satellites and rockets, determine the energy separation state during ground testing that meets the requirements for experimental authenticity. That is, analyze the difference between space and ground conditions under ideal conditions through step S104. The difference between space and ground conditions refers to the difference between the satellite and rocket separation test specimen in flight and during ground testing. In specific implementation, the reference moment is when the separation spring rod of the satellite locking separation device disengages from the lower port of the satellite docking frame (i.e., the interaction force between the launch vehicle and the satellite is zero). Ideal conditions refer to the absence of jamming in the satellite locking separation device compared to unexpected conditions. Unexpected conditions, on the other hand, refer to jamming of the locking device.

[0059] The relative separation velocity of the first satellite and rocket in this embodiment is calculated based on the following formula:

[0060]

[0061] The relative separation velocity of the second satellite and rocket is calculated based on the following formula:

[0062]

[0063] The relative separation velocity of the third satellite and rocket is calculated based on the following formula:

[0064]

[0065] Wherein, m1 is the mass of the satellite, m2 is the mass of the launch satellite support, m3 is the mass of the final stage of the launch vehicle, F0 is the initial separation force of the separation spring, h is the extension height of the separation spring of the launch satellite support, and g is the acceleration due to gravity.

[0066] Furthermore, the energy separation state of the ground test process that meets the requirements of test authenticity is determined by comparing the relative separation velocities of the first satellite and rocket, the second satellite and rocket, and the third satellite and rocket.

[0067] Based on the specific application examples above, substituting the corresponding parameters from Tables 1 and 2 into the formula, we can calculate the relative separation velocity Δv1 of the first satellite and rocket in flight mode as 0.66 m / s; the relative separation velocity Δv2 of the second satellite and rocket in the ground test process with energy enabled as separation occurs as 4.6 m / s; and the relative separation velocity Δv3 of the third satellite and rocket in the ground test process with energy disabled as separation occurs as 1.7 m / s. By comparison, we can see that the relative separation velocity Δv3 of the third satellite and rocket is closer to the relative separation velocity Δv1 of the first satellite and rocket. Therefore, for the sake of experimental authenticity, the energy disabled state during the ground test process should be selected.

[0068] S105: Based on the geometric and kinematic parameters, analyze and calculate the relative rotational angular velocity of the first satellite-rocket in flight, the relative rotational angular velocity of the second satellite-rocket in the energy-enabled state during ground testing, and the relative rotational angular velocity of the third satellite-rocket in the energy-disabled state during ground testing under unexpected conditions; by comparing the relative rotational angular velocity of the second satellite-rocket with that of the third satellite-rocket, determine the energy-disconnection state during ground testing that meets the safety requirements of the test; that is, analyze the difference between the space and ground under unexpected conditions through step S105, where the difference between the space and ground refers to the difference between the satellite-rocket separation test specimen in flight and during ground testing.

[0069] The relative rotational angular velocity of the first star-rocket in this embodiment is calculated based on the following formula:

[0070]

[0071] The relative rotational angular velocity of the second satellite rocket is calculated based on the following formula:

[0072]

[0073] The relative rotational angular velocity of the third satellite is calculated based on the following formula:

[0074]

[0075] Wherein, I1 is the moment of inertia of the satellite, I2 is the moment of inertia of the launch satellite support, I3 is the moment of inertia of the final stage of the launch vehicle, F0 is the initial separation force of the separation spring, m2 is the mass of the launch satellite support, h is the extension height of the separation spring of the launch satellite support, and g is the acceleration due to gravity.

[0076] Furthermore, the energy separation state of the ground test process that meets the test safety requirements is determined by comparing the relative rotational angular velocities of the second satellite and the third satellite.

[0077] Based on the specific application examples above, substituting the corresponding parameters from Tables 1 and 2 into the formula, we can calculate the relative rotational angular velocity of the first satellite and rocket in flight state as Δω1 = 0.15 rad / s; the relative rotational angular velocity of the second satellite and rocket in the ground test process with energy separation enabled as Δω2 = 6.16 rad / s; and the relative rotational angular velocity of the third satellite and rocket in the ground test process with energy separation disabled as Δω3 = 1.63 rad / s. By comparison, it can be seen that the relative rotational angular velocity of the third satellite and rocket as Δω3 is closer to that of the first satellite and rocket as Δω1. Therefore, for safety reasons, energy separation must be disabled during the ground test.

[0078] S106: Based on the geometric model, analyze the positional relationship between the satellite and the satellite support during the separation process under unexpected operating conditions, and obtain the minimum distance between the satellite and the satellite support based on the positional relationship; Step S106 of this embodiment further includes: determining the first stage and the second stage of the separation process between the satellite and the satellite support under unexpected operating conditions; wherein, the first stage is to take the side of the satellite support connected to the satellite as the first fixed end, and under the action of gravity, the satellite support undergoes rigid body rotation around the first fixed end until the lower edge of the satellite support contacts the ground protective pad; the second stage is to take the contact point between the lower edge of the satellite support and the ground protective pad as the second fixed end, the first fixed end is released from constraint, and under the action of gravity, the satellite support undergoes rigid body rotation around the second fixed end until the satellite support is in complete contact with the ground protective pad; based on the positional relationship between the satellite and the satellite support in the first stage and the second stage of the geometric model, extract the minimum distance between the satellite and the satellite support. In this embodiment, step S106 analyzes the geometric interference of unexpected ground test conditions. Specifically, based on the geometric model constructed in step S103, it analyzes the relationship between the satellite and satellite support separation process under unexpected conditions (jamming on one side of the locking device). Under unexpected conditions, the separation motion of the satellite and satellite support is divided into two stages. First, taking the side of the satellite support connected to the satellite as the first fixed end, the satellite support undergoes rigid body rotation around the first fixed end under the action of gravity until the lower edge of the satellite support contacts the ground protective pad. Then, taking the contact point between the lower edge of the satellite support and the ground protective pad as the second fixed end, the first fixed end is released from constraint, and the satellite support undergoes rigid body rotation around the second fixed end under the action of gravity until the satellite support is in complete contact with the ground protective pad. Based on the analysis of the above positional relationship using the geometric model, this embodiment can extract the minimum distance d between the satellite and satellite support in the two motion stages. min In specific implementation, taking an initial height of 600mm between the satellite support and the ground protective pad as an example, the minimum distance between the first and second movement phases is as follows: Figure 4 d min =7.88mm.

[0079] S107: If the minimum distance is less than the safe distance for satellite-rocket separation, adjust the height of the lower edge of the satellite support from the ground protection pad, and return to recalculate the minimum distance until the minimum distance is greater than the safe distance for satellite-rocket separation. That is, if d min < S, adjust the height D of the lower edge of the satellite support from the ground protection pad, and return to loop and execute steps S103 - S106 until the minimum distance is greater than or equal to the separation safe distance. As Figure 5 shown, when the height of the satellite support and the ground protection pad is 900 mm, the minimum distance d min = 248.51 mm in the first motion stage and the second motion stage, meeting the requirement of not being less than the separation safe distance.

[0080] Subsequently, determine the ground test separation energy state according to steps S104 - S105; and determine the height of the satellite support and the ground protection pad according to step S107, finally completing the safety design of the satellite-rocket separation test site; in combination with a specific application example, the ground test separation energy can be determined to be in a prohibited state according to steps S104 - S105; determine the height of the satellite support and the ground protection pad to be 900 mm according to step S107, and finally complete the near-field safety design of the satellite-rocket separation test.

[0081] See Figure 2 , the specific implementation process of the design method for the near-field safety of the satellite-rocket separation test in the present invention is: determine the relevant parameters of the satellite-rocket separation test specimens, and the relevant parameters are geometric parameters and corresponding kinematic parameters; establish a geometric model and analyze the differences between space and ground in ideal conditions and the differences between space and ground in unanticipated conditions respectively; determine the ground test separation energy state based on the test authenticity for the differences between space and ground in ideal conditions; determine the ground test separation energy state based on the test safety for the differences between space and ground in unanticipated conditions; perform geometric interference analysis on the unanticipated conditions of the ground test according to the geometric model, extract the minimum distance between the satellite and the satellite support during the movement process, and judge whether the minimum distance is less than the safe distance for satellite-rocket separation. If it is less, adjust the height of the satellite support and the ground protection pad until the minimum distance is not less than the safe distance for satellite-rocket separation, and then determine the adjusted height of the satellite support and the ground protection pad; finally, output the determined ground test separation energy state and the height of the satellite support and the ground protection pad, completing the near-field safety design of the satellite-rocket separation test.

[0082] The present invention also provides a storage medium for storing such as Figure 1The computer program for the design method of near-field safety in the star-rocket separation test. For example, computer program instructions, when executed by a computer, can invoke or provide the methods and / or technical solutions according to this application through the operation of the computer. The program instructions invoking the methods of this application may be stored in a fixed or removable storage medium, and / or transmitted via data streams in broadcast or other signal carrying media, and / or stored in the storage medium of a computer device operating according to the program instructions. Here, one embodiment according to this application includes, as... Figure 1 The computer device for designing a near-field safety method for a satellite-rocket separation test, as shown, preferably includes a storage medium for storing a computer program and a processor for executing the computer program, wherein when the computer program is executed by the processor, the computer device is triggered to execute the method and / or technical solution based on the foregoing embodiments.

[0083] In summary, the design method for near-field safety of spacecraft separation tests described in this invention can rapidly assess near-field safety under complex boundary and unexpected operating conditions based on motion geometry interference analysis and rigid body dynamics equations, saving manpower and resources and reducing test risks.

[0084] Of course, the present invention may have other various embodiments. Without departing from the spirit and essence of the present invention, those skilled in the art can make various corresponding changes and modifications according to the present invention, but these corresponding changes and modifications should all fall within the protection scope of the appended claims.

Claims

1. A design method for near-field safety of a satellite-rocket separation test, characterized by, The method comprises the following steps: Obtaining geometric parameters of a star-rocket separation test sample; wherein the geometric parameters at least include a star-rocket docking frame radius, a satellite device envelope size of a satellite device penetrating into a satellite support, a launch satellite support separation spring extension height, a satellite support height, a satellite support lower edge radius, a satellite support lower edge distance from a ground pad, and a star-rocket separation safety distance; Obtaining kinematic parameters corresponding to the star-rocket separation test sample; wherein the kinematic parameters include a satellite mass, a satellite rotational inertia, a launch satellite support mass, a launch satellite support rotational inertia, a launch final stage mass, a launch final stage rotational inertia, and a separation spring initial separation resultant force; Establishing a geometric model of the satellite and the launch satellite support according to the geometric parameters; According to the geometric parameters and the kinematic parameters, analyzing and calculating a first star-rocket relative separation speed of the star-rocket separation test sample in a flight state under an ideal working condition, a second star-rocket relative separation speed in a separation energy enabled state during a ground test process, and a third star-rocket relative separation speed in a separation energy prohibited state during the ground test process; according to a comparison between the first star-rocket relative separation speed and the second star-rocket relative separation speed and the third star-rocket relative separation speed, determining a separation energy state during the ground test process that meets a test authenticity requirement; According to the geometric parameters and the kinematic parameters, analyzing and calculating a first star-rocket relative rotational angular velocity of the star-rocket separation test sample in the flight state under an unexpected working condition, a second star-rocket relative rotational angular velocity in the separation energy enabled state during the ground test process, and a third star-rocket relative rotational angular velocity in the separation energy prohibited state during the ground test process; through a comparison between the second star-rocket relative rotational angular velocity and the third star-rocket relative rotational angular velocity, determining a separation energy state during the ground test process that meets a test safety requirement; According to the geometric model, analyzing a positional relationship of a satellite and a satellite support separation process under an unexpected working condition, and obtaining a minimum distance between the satellite and the satellite support based on the positional relationship; If the minimum distance is less than the star-rocket separation safety distance, adjusting a satellite support lower edge distance from a ground pad, and returning to recalculate the minimum distance until the minimum distance is greater than the star-rocket separation safety distance.

2. The method of designing the near-field safety of satellite-rocket separation tests according to claim 1, characterized in that, The first star-rocket relative separation speed is calculated based on the following formula: The second star-rocket relative separation speed is calculated based on the following formula: The third star-rocket relative separation speed is calculated based on the following formula: Wherein, m1 is the satellite mass, m2 is the launch satellite support mass, m3 is the launch final stage mass, F0 is the separation spring initial separation resultant force, h is the launch satellite support separation spring extension height, and g is the gravitational acceleration.

3. The method of designing the near-field safety of satellite-rocket separation tests according to claim 1, characterized in that, The first star-rocket relative rotational angular velocity is calculated based on the following formula: The second star-rocket relative rotational angular velocity is calculated based on the following formula: The third star-rocket relative rotational angular velocity is calculated based on the following formula: Wherein, I1 is the satellite moment of inertia, I2 is the satellite carrier moment of inertia, I3 is the carrier final stage moment of inertia, F0 is the initial separation spring combined force, m2 is the satellite carrier mass, h is the satellite carrier separation spring extension height, and g is the gravity acceleration.

4. The method of designing the near-field safety of satellite-rocket separation tests according to claim 1, characterized in that, The position relationship of the satellite and the satellite carrier during the separation process under the unexpected working condition is analyzed according to the geometric model, and the minimum distance between the satellite and the satellite carrier is obtained based on the position relationship. The first stage and the second stage of the satellite and the satellite carrier during the separation process under the unexpected working condition are determined, wherein the first stage is that the satellite carrier rotates as a rigid body around the first fixed end under the action of gravity until the satellite carrier contacts the ground protection pad, and the second stage is that the satellite carrier rotates as a rigid body around the second fixed end under the action of gravity until the satellite carrier completely contacts the ground protection pad. The minimum distance between the satellite and the satellite carrier is extracted according to the position relationship of the satellite and the satellite carrier in the first stage and the second stage.

5. A storage medium, characterized by A computer program for executing the design method of the near-field safety of the satellite-rocket separation test in any one of claims 1-4.

6. A computer device comprising a storage medium, a processor, and a computer program stored on the storage medium and executable on the processor, characterized in that, The processor executes the computer program to realize the design method of the near-field safety of the satellite-rocket separation test in any one of claims 1-4.

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