A method for modeling the dynamics of a spinning spacecraft with mass jump characteristics

Abstract

The application discloses a kind of spin spacecraft dynamics modeling methods with mass mutation characteristics, comprising the following steps: step 100, according to the structural characteristics of whole satellite system, it is decomposed into satellite platform body, three-axis superconducting magnet system, rotating separation device, momentum exchange device and separable load.Step 200, based on the orbit and structural characteristics of satellite, the derivation assumption about the whole satellite dynamics equation is proposed to build ideal modeling environment.The dynamics modeling method of part of spin spacecraft in the application considers mass mutation, which can be used not only for on-orbit delivery spacecraft, but also for other spacecraft with mass mutation characteristics, and the vector mechanics method is used in the modeling process of the application, which can more intuitively reflect the system dynamics state, and is beneficial to in-depth study of satellite nutation mechanism from the overall perspective.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and more specifically to a method for modeling the dynamics of a spin spacecraft with mass mutation characteristics. Background Technology

[0002] With the increasing demand for human exploration and utilization of near-Earth space, traditional single-spin and dual-spin spacecraft are unable to achieve their intended service goals, such as continuous multi-payload delivery and high-speed on-orbit delivery. Therefore, some spin spacecraft have emerged.

[0003] Some spinning spacecraft possess significant advantages such as flexibility, controllability, and stability, enabling them to conduct more complex or specialized on-orbit missions, thus highlighting their growing importance in space resource utilization. Typical examples include NASA's proposed space target transport based on a space-based rotational momentum exchange tether, and the low-Earth orbit in-situ geomagnetic energy storage delivery based on the space-based rotational acceleration principle proposed by the Institute of Mechanics, Chinese Academy of Sciences (patents 201900773631.X and 201910774225.5). These partially spinning spacecraft often exhibit abrupt mass changes, posing a significant challenge to the dynamic modeling of the entire system.

[0004] Currently, available literature and authorized patents (represented by patent publication number CN109033604A) mainly focus on satellites with rotating loads or elastic / flexible appendages, providing system dynamics modeling methods. However, the dynamic equations can only study the force conditions of local individual components, and the modeling does not consider the mass mutation process. Therefore, it is impossible to study the nutation instability mechanism and characteristics of the satellite platform caused by this feature. Without understanding the nutation law, it is impossible to propose effective active / passive suppression control laws and design controllers. Therefore, the dynamic equations of partially spinning spacecraft with mass mutation characteristics are crucial, and modeling methods urgently need research. This work will lay a theoretical foundation for the development of special spacecraft dynamics and control. Summary of the Invention

[0005] The purpose of this invention is to provide a modeling method for the dynamics of a spin spacecraft with mass mutation characteristics, so as to solve the technical problem that the modeling method for the dynamic equations of some spin spacecraft with mass mutation characteristics urgently needs to be studied in the prior art.

[0006] To solve the above-mentioned technical problems, the present invention specifically provides the following technical solution:

[0007] A method for modeling the dynamics of a spin spacecraft with mass mutation characteristics includes the following steps:

[0008] Step 100: Based on the structural characteristics of the entire satellite system, it is decomposed into the satellite platform body, the three-axis superconducting magnet system, the rotation separation device, the momentum exchange device, and the separable payload;

[0009] Step 200: Based on the satellite's orbital and structural characteristics, propose derivation assumptions about the whole-satellite dynamics equations to construct an ideal modeling environment;

[0010] Step 300: To facilitate the description of each component, the following coordinate system is established:

[0011] The system's centroid orbital coordinate system OXYZ is established with the system's centroid as its origin.

[0012] ObXbYbZb, with the center of mass of the satellite platform as the origin;

[0013] OcXcYcZc, with the center of mass of the rotating separation device as the origin;

[0014] OdXdYdZd, with the separable load centroid as the origin;

[0015] Step 400: Based on the derivation assumptions of Step 200, the system after load separation is considered as a two-rigid-body system. The system's angular momentum is expressed in four parts: the angular momentum generated by the rotating separation device around its own center of mass, the angular momentum generated by the center of mass of the rotating separation device around the system's combined center of mass, the angular momentum generated by the satellite platform's rotation around its own center of mass, and the angular momentum generated by the satellite platform's center of mass around the system's combined center of mass. These are defined as follows: and respectively calculate;

[0016] Calculate the resultant angular momentum of the four systems mentioned above;

[0017] Based on the attitude transformation matrix between the reference frame of the rotating separation device and the reference frame of the satellite platform, the resultant angular momentum is transformed to obtain the resultant angular momentum of the system relative to the center of mass.

[0018] Step 500: Calculate the net angular momentum of the system relative to the center of mass obtained in step 400 using the angular momentum theorem, and obtain the overall dynamic equation of the satellite system by combining the derivation assumptions.

[0019] As a preferred embodiment of the present invention, in step 100;

[0020] In step 100;

[0021] The satellite platform body consists of a rigid outer shell, internal rigid components, and a momentum flywheel.

[0022] The momentum exchange device consists of a stator and a rotor, and the stator is fixedly connected to the satellite platform body;

[0023] The triaxial superconducting magnet system is fixedly connected to the satellite platform body through the stator, and the triaxial superconducting magnet system can influence the attitude of the satellite platform body through external torque;

[0024] The rotary separation device is fixedly installed on the rotor, and a slide rail for continuously conveying the separable load is provided on the rotary separation device, and a controllable instantaneous release mechanism is provided at the end of the slide rail;

[0025] The separable load is a single rigid body that is slidably mounted on the slide rail, and the separable load is launched through the controllable instantaneous release mechanism.

[0026] As a preferred embodiment of the present invention, the following assumptions are derived in step 200:

[0027] In dynamic modeling and analysis, the satellite platform body, the mass separation rotating separation device, and the superconducting magnet system are all assumed to be rigid bodies.

[0028] Neglecting the mass of the momentum exchange device and the effects of the gap and damping between the stator and rotor;

[0029] Assume that the mass separation rotating device and the load to be separated separate instantaneously without any impact between them;

[0030] Assume the satellite was in a stable state before separation.

[0031] As a preferred embodiment of the present invention, the specific steps for establishing multiple coordinate systems in step 300 are as follows:

[0032] The system's centroid orbital coordinate system OXYZ: The origin O is at the system's centroid, the OZ axis points to the Earth's center, the OX axis is in the satellite's orbital plane, perpendicular to the OZ axis and pointing in the direction of the satellite's motion, and the OY axis, together with the OX and OZ axes, forms a right-handed coordinate system;

[0033] The satellite platform's body coordinate system ObXbYbZb: origin O b Located at the center of mass of the satellite platform, when the three azimuth angles of the platform relative to the orbital coordinate system are zero, the direction of each axis is consistent with the direction of each axis of the orbital coordinate system OXYZ.

[0034] The coordinate system of the rotating separation device is OcXcYcZc: the origin of the coordinate system is O c Located at the intersection of the rotating axis and the rotating plane of the rotating separation device, and assuming that at the initial moment, the axes of the rotating separation device point in the same direction as the coordinate axes of the platform body coordinate system;

[0035] The separable load body coordinate system OdXdYdZd: the origin of the coordinate system O d Located at the center of mass of the separable load 5, the three coordinate axes before separation are parallel to the corresponding coordinate axes of the body coordinate system of the rotating separation device.

[0036] In a preferred embodiment of the present invention, in step 400, after the separable load is separated, the angular momentum generated by the rotating separation device about its own center of mass is:

[0037]

[0038] in; The moment of inertia of the rotating separation device and the delivered load relative to the center of mass before separation;

[0039] ω is the relative angular velocity between the satellite platform's body reference frame and the center-of-mass orbit reference frame;

[0040] Ω represents the relative angular velocity between the optional device reference frame and the satellite platform's own frame.

[0041] As a preferred embodiment of the present invention, in step 400, the centroid of the rotating separation device revolves around the system.

[0042] The angular momentum generated by the combined center of mass is:

[0043]

[0044] Where, m c For the mass of the rotary separator;

[0045] r c Let be the radius vector of the center of mass of the rotating separation device relative to the combined center of mass of the system;

[0046] ω is the relative angular velocity between the satellite platform's body reference frame and the center-of-mass orbit reference frame;

[0047] Ω represents the relative angular velocity between the optional device reference frame and the satellite platform's own frame.

[0048] As a preferred embodiment of the present invention, in step 400, the angular momentum generated by the satellite platform body rotating around its own center of mass is:

[0049]

[0050] in, The moment of inertia of the satellite platform and the superconducting device relative to the combined center of mass;

[0051] ω is the relative angular velocity between the satellite platform's body reference frame and the center-of-mass orbit reference frame.

[0052] As a preferred embodiment of the present invention, in step 400, the platform's center of mass revolves around the system's center of mass production.

[0053] Angular momentum of life:

[0054]

[0055] Where, mb The combined mass of the satellite platform and the superconducting device;

[0056] r b Let be the radius vector of the satellite platform's center of mass relative to the system's combined center of mass;

[0057] ω is the relative angular velocity between the satellite platform's body reference frame and the center-of-mass orbit reference frame;

[0058] Ω represents the relative angular velocity between the optional device reference frame and the satellite platform's own frame.

[0059] As a preferred embodiment of the present invention, in step 400, the calculation time of the moment of inertia of the center of mass of the rotating separation device relative to the system's combined center of mass is the instant when the separable load separates from the rotating separation device, and it is assumed that there is no impact between the separable load and the rotating separation device.

[0060] Compared with the prior art, the present invention has the following advantages:

[0061] The dynamic modeling method for some spinning spacecraft in this invention considers mass mutation. This modeling method can be used not only for spacecraft deployed in orbit, but also for other spacecraft with mass mutation characteristics. Furthermore, this invention adopts vector mechanics in the modeling process, which can more intuitively reflect the dynamic state of the system and is conducive to in-depth study of satellite nutation mechanism from a holistic perspective. Attached Figure Description

[0062] To more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings in the following description are merely exemplary, and those skilled in the art can derive other embodiments based on the provided drawings without creative effort.

[0063] Figure 1 A schematic diagram of the structure of a spin spacecraft dynamics modeling method with mass mutation characteristics is provided for embodiments of the present invention;

[0064] The labels in the diagram represent the following:

[0065] 1. Satellite platform body; 2. Triaxial superconducting magnet system;

[0066] 3. Rotary separation device, 301 slide rail;

[0067] 4. Momentum exchange device, 401 stator, 402 rotor;

[0068] 5. Separable load. Detailed Implementation

[0069] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0070] like Figure 1 As shown, this invention provides a method for modeling the dynamics of a spin spacecraft with mass mutation characteristics, comprising the following steps:

[0071] Step 100: Based on the structural characteristics of the entire satellite system, it is decomposed into the satellite platform body 1, the three-axis superconducting magnet system 2, the rotation separation device 3, the momentum exchange device 4, and the separable payload 5.

[0072] Among them, the satellite platform body 1 consists of a rigid outer shell, internal rigid components, and a momentum flywheel.

[0073] The momentum exchange device 4 consists of a stator 401 and a rotor 402, with the stator 401 fixedly connected to the satellite platform body 1. The relative rotation direction and speed of the stator 401 and the rotor 402 are controlled by the system.

[0074] The three-axis superconducting magnet system is fixedly connected to the satellite platform body 1 via stator 401. The three-axis superconducting magnet system can obtain momentum torque by interacting with the geomagnetic field to inject energy into the system, and can also obtain external torque to adjust the platform attitude or unload the flywheel.

[0075] The rotary separation device 3 is fixedly installed on the rotor 402. A slide rail 301 for continuously conveying the separable load 5 is provided on the rotary separation device 3, and a controllable instantaneous release mechanism is provided at the end of the slide rail 301.

[0076] The separable load 5 is a single rigid body that is slidably mounted on the slide rail 301. The separable load 5 is launched through a controllable instantaneous release mechanism.

[0077] Step 200: Based on the satellite's orbital and structural characteristics, propose assumptions for the derivation of the whole-satellite dynamics equations;

[0078] 1. In the dynamic modeling and analysis, the satellite platform body 1, the rotating separation device 3, and the triaxial superconducting magnet system 2 are all assumed to be rigid bodies.

[0079] 2. The mass of the momentum exchange device 4, and the effects of the gap and damping between the stator 401 and the rotor 402 are ignored.

[0080] 3. Assume that the rotating separation device 3 and the separable load 5 separate instantaneously without impact between them.

[0081] 4. Assume the satellite is in a stable state before separation.

[0082] Step 300: To facilitate the description of each component, the following coordinate system is introduced:

[0083] The system's centroid orbital coordinate system OXYZ: The origin O is at the system's centroid, the OZ axis points to the Earth's center, the OX axis is in the satellite's orbital plane, perpendicular to the OZ axis and pointing in the direction of the satellite's motion, and the OY axis, together with the OX and OZ axes, forms a right-handed coordinate system;

[0084] Satellite platform body 1 coordinate system ObXbYbZb: The origin Ob is located at the center of mass of satellite platform body 1. When the three azimuth angles of the platform relative to the orbit coordinate system are zero, the direction of each axis is consistent with the direction of each axis of the orbit coordinate system OXYZ.

[0085] The coordinate system OcXcYcZc of the rotating separation device 3: The origin Oc of the coordinate system is located at the intersection of the rotating axis and the rotating plane of the rotating separation device 3, and it is assumed that at the initial moment, the axes of the rotating separation device 3 are aligned with the coordinate axes of the platform body coordinate system.

[0086] The coordinate system OdXdYdZd of the separable load 5: The origin Od of the coordinate system is located at the centroid of the separable load 5. Before separation, the three coordinate axes are parallel to the corresponding coordinate axes of the body coordinate system of the rotating separation device 3.

[0087] Step 400, based on the assumptions derived in Step 200, treats the system after load separation as a two-rigid-body system (Basic Assumption 1 of Step 200). The system's angular momentum is expressed in four parts: the angular momentum generated by the rotating separation device about its own center of mass, the angular momentum generated by the center of mass of the rotating separation device about the system's combined center of mass, the angular momentum generated by the platform rotating about its own center of mass, and the angular momentum generated by the platform's center of mass about the system's combined center of mass. These are defined as follows:

[0088] The following are the meanings of the symbols in the formulas below:

[0089] Symbol meaning

[0090] d Distance between the load and the center of mass of the rotating separation device

[0091] m c Mass of rotary separator

[0092] m b Combined mass of satellite platform and superconducting device

[0093] Δm Mass of the load to be separated

[0094] h Distance between the rotating plane of the optional device and the center of gravity of the platform

[0095] m1 Mass of load to be separated

[0096] m2 Mass of load to be separated

[0097] e Mass of load to be separated

[0098] u Mass of load to be separated

[0099] ω: Relative angular velocity between the satellite platform's body reference frame and the center-of-mass orbital reference frame.

[0100] Ω Relative angular velocity between the optional device reference frame and the satellite platform's own frame r c The radius vector of the center of mass of the rotating separation device relative to the combined center of mass of the system

[0101] r b The radius vector of the satellite platform's center of mass relative to the system's combined center of mass

[0102] Separation load and rotational inertia of the rotating separation device relative to the center of mass

[0103] Moment of inertia of the rotating separation device relative to its own center of mass

[0104] Moment of inertia of the satellite platform and superconducting device relative to the combined center of mass

[0105] Moment of inertia of the load relative to the center of mass of the optional device

[0106] Moment of inertia of the center of mass of the rotating separation device relative to the combined center of mass of the system

[0107] Attitude transformation matrix from C-rotation separation device reference frame to platform body reference frame

[0108] The control torque generated by the interaction of the L superconducting magnetic moment with the Earth's magnetic field

[0109] Among them, 1. The calculation process is as follows:

[0110] Let the inertia matrix of the delivered payload relative to its own center of mass be... If the distance between the load and the center of mass c of the rotating separation device is d, then the moment of inertia of the load relative to the center of mass c of the rotating separation device in the reference frame of the rotating separation device can be expressed as:

[0111] in,

[0112] Formula (1)

[0113] For a rotating separation device, the centroid deviation due to load separation is:

[0114] Formula (2)

[0115] The moment of inertia of the rotating separator after mass separation can be expressed as:

[0116] Formula (3)

[0117] in, The moment of inertia of the rotating separation device and the delivered load relative to the center of mass before separation.

[0118] After the load is separated, the angular momentum generated by the rotating separation device about its own center of mass is:

[0119] Formula (4)

[0120] Similarly, the calculation results for the angular momentum generated by the center of mass of the rotating separation device around the system's combined center of mass, the angular momentum generated by the platform rotating around its own center of mass, and the angular momentum generated by the platform's center of mass around the system's combined center of mass are as follows:

[0121] 2. Calculation

[0122] Angular momentum generated by the center of mass of the rotating separation device around the system's combined center of mass:

[0123] Formula (5)

[0124] 3. Calculation

[0125] Angular momentum generated by the platform rotating around its own center of mass:

[0126] Formula (6)

[0127] 4. Calculation

[0128] Angular momentum generated by the platform's center of mass around the system's combined center of mass:

[0129] Formula (7)

[0130] 5. Calculation and Simplification of Net Angular Momentum

[0131] The resultant angular momentum can be expressed as: Formula (8)

[0132] in,

[0133]

[0134] Let be the moment of inertia of the center of mass of the rotating component relative to the net center of mass at t=0;

[0135] C is the attitude transformation matrix between the reference frame of the rotating separation device and the reference frame of the platform body.

[0136] Since load separation occurs instantaneously, there is no impact between the load and the rotating separation device (basic assumption 3 of step 200), and also:

[0137] Formula (9)

[0138] Where: ΔH is the angular momentum loss due to load separation, expressed as:

[0139]

[0140] The initial kinetic energy of the system after load separation is:

[0141] Formula (10) E(t=0)=E0-ΔE;

[0142] This includes the mechanical energy loss due to load separation.

[0143]

[0144] Step 500: Calculate the net angular momentum of the system relative to the center of mass obtained from formula (9) using the angular momentum theorem, and obtain:

[0145] Formula (11)

[0146]

[0147] Where M is the sum of the disturbance torque and the control torque experienced by the satellite.

[0148] set up Simplifying, we get:

[0149] Formula (12)

[0150] This equation is the overall dynamic equation of the satellite system.

[0151] The above embodiments are merely exemplary embodiments of this application and are not intended to limit this application. The scope of protection of this application is defined by the claims. Those skilled in the art can make various modifications or equivalent substitutions to this application within its substance and scope of protection, and such modifications or equivalent substitutions should also be considered to fall within the scope of protection of this application.

Claims

1. A method for modeling the dynamics of a spin spacecraft with mass mutation characteristics, characterized in that, Includes the following steps: Step 100: Based on the structural characteristics of the whole satellite system, it is decomposed into the satellite platform body (1), the three-axis superconducting magnet system (2), the rotating separation device (3), the momentum exchange device (4), and the separable payload (5); The satellite platform body (1) consists of a rigid outer shell, internal rigid components, and a momentum flywheel; The momentum exchange device (4) consists of a stator (401) and a rotor (402), and the stator (401) is fixedly connected to the satellite platform body (1); The triaxial superconducting magnet system is fixedly connected to the satellite platform body (1) through the stator (401), and the triaxial superconducting magnet system can adjust the attitude of the satellite platform body (1) by external torque; The rotary separation device (3) is fixedly installed on the rotor (402), and a slide rail (301) for continuously conveying the separable load (5) is provided on the rotary separation device (3), and a controllable instantaneous release mechanism is provided at the end of the slide rail (301); The separable load (5) is a single rigid body that is slidably mounted on the slide rail (301), and the separable load (5) is launched through the controllable instantaneous release mechanism; Step 200: Based on the satellite's orbital and structural characteristics, propose derivation assumptions for the whole-satellite dynamics equations to construct an ideal modeling environment. The derivation assumptions are as follows: In the dynamic modeling and analysis, the satellite platform body (1), the rotating separation device (3), and the superconducting magnet system (2) are all assumed to be rigid bodies; Ignore the mass of the momentum exchange device (4), and the effects of the gap and damping between the stator (401) and the rotor (402); Assuming that the rotating separation device (3) and the load to be separated separate instantaneously without any impact between them; Assume the satellite was in a stable state before separation; Step 300: To facilitate the description of each component, the following coordinate system is established: The system's centroid orbital coordinate system OXYZ is established with the system's centroid as its origin. The origin is ObXbYbZb, with the center of mass of the satellite platform body (1) as the origin; OcXcYcZc, with the centroid of the rotating separation device (3) as the origin; OdXdYdZd with the centroid of the separable load (5) as the origin; Step 400: Based on the derivation assumptions of Step 200, the system after load separation is considered as a two-rigid-body system. The angular momentum of the system is expressed in four parts: the angular momentum generated by the rotating separation device (3) around its own center of mass, the angular momentum generated by the center of mass of the rotating separation device (3) around the combined center of mass of the system, the angular momentum generated by the satellite platform body (1) rotating around its own center of mass, and the angular momentum generated by the center of mass of the satellite platform body (1) around the combined center of mass of the system. These are defined as follows: and respectively calculate; Calculate the resultant angular momentum of the four systems mentioned above; The resultant angular momentum is transformed based on the attitude transformation matrix between the reference frame of the rotating separation device (3) and the reference frame of the satellite platform body (1) to obtain the resultant angular momentum of the system relative to the center of mass; Step 500: Calculate the net angular momentum of the system relative to the center of mass obtained in step 400 using the angular momentum theorem, and obtain the overall dynamic equation of the satellite system by combining the derivation assumptions.

2. The method for modeling the dynamics of a spin spacecraft with a mass mutation characteristic according to claim 1, characterized in that, The specific steps for establishing multiple coordinate systems in step 300 are as follows: The system's centroid orbital coordinate system OXYZ: The origin O is at the system's centroid, the OZ axis points to the Earth's center, the OX axis is in the satellite's orbital plane, perpendicular to the OZ axis and pointing in the direction of the satellite's motion, and the OY axis, together with the OX and OZ axes, forms a right-handed coordinate system; The coordinate system of the satellite platform body (1) is ObXbYbZb: origin O b Located at the center of mass of the satellite platform body (1), when the three azimuth angles of the platform relative to the orbital coordinate system are zero, the direction of each axis is consistent with the direction of each axis of the orbital coordinate system OXYZ. The rotating separation device (3) has a body coordinate system OcXcYcZc: the origin of the coordinate system is O c Located at the intersection of the rotating axis and the rotating plane of the rotating separation device (3), and assuming that at the initial moment, each axis of the rotating separation device (3) points in the same direction as each coordinate axis of the platform body coordinate system; The separable load (5) body coordinate system OdXdYdZd: coordinate system origin O d Located at the centroid of the separable load (5), the three coordinate axes before separation are parallel to the corresponding coordinate axes of the body coordinate system of the rotating separation device (3).

3. The method for modeling the dynamics of a spin spacecraft with a mass mutation characteristic according to claim 2, characterized in that, In step 400, after the separable load (5) is separated, the angular momentum generated by the rotating separation device (3) about its own center of mass is: in; The moment of inertia of the rotating separation device (3) and the delivery load relative to the center of mass before separation; ω is the relative angular velocity between the satellite platform's body reference frame and the center-of-mass orbit reference frame; Ω represents the relative angular velocity between the optional device reference frame and the satellite platform's own frame.

4. The method for modeling the dynamics of a spin spacecraft with a mass mutation characteristic according to claim 2, characterized in that, In step 400, the angular momentum generated by the center of mass of the rotating separation device (3) around the combined center of mass of the system is: Where, m c For the mass of the rotary separator; r c The radius vector of the center of mass of the rotating separation device (3) relative to the combined center of mass of the system; ω is the relative angular velocity between the satellite platform body (1) reference frame and the center of mass orbit reference frame; Ω represents the relative angular velocity between the optional device reference frame and the satellite platform body (1) frame.

5. The method for modeling the dynamics of a spin spacecraft with a mass mutation characteristic according to claim 2, characterized in that, In step 400, the angular momentum generated by the satellite platform body (1) rotating around its own center of mass is: in, The moment of inertia of the satellite platform body (1) and the superconducting device relative to the center of mass; ω is the relative angular velocity between the satellite platform body (1) reference frame and the center of mass orbit reference frame.

6. The method for modeling the dynamics of a spin spacecraft with a mass mutation characteristic according to claim 2, characterized in that, In step 400, the angular momentum generated by the platform's center of mass around the system's combined center of mass is: Where, m b The combined mass of the satellite platform body (1) and the superconducting device; r b The radius vector of the center of mass of the satellite platform (1) relative to the combined center of mass of the system; ω is the relative angular velocity between the satellite platform body (1) reference frame and the center of mass orbit reference frame; Ω represents the relative angular velocity between the optional device reference frame and the satellite platform body (1) frame.

7. The method for modeling the dynamics of a spin spacecraft with a mass mutation characteristic according to claim 1, characterized in that, In step 400, the moment of inertia of the center of mass of the rotating separation device (3) relative to the system's combined center of mass is calculated at the instant when the separable load (5) separates from the rotating separation device (3), and it is assumed that there is no impact between the separable load (5) and the rotating separation device (3).

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