A nutation characteristics analysis method and active suppression system for spinning spacecraft

Abstract

The present invention discloses a method for analyzing the nutation characteristics of a spinning spacecraft and an active suppression system, comprising the following steps: Step 100: Decomposing the entire satellite system into a satellite platform body, a three-axis superconducting magnet system, a rotation separation device, a momentum exchange device, and a detachable payload based on its structural characteristics. Step 200: Proposing derivation hypotheses for the dynamic equations of the entire satellite based on the satellite's orbit and structural characteristics to construct an ideal modeling environment. The present invention describes a method for analyzing the nutation characteristics of a partially spinning spacecraft, incorporating mass mutations. This analysis method can be used not only for on-orbit delivery spacecraft but also for other spacecraft with mass mutation characteristics. The control method utilizes a fuzzy PID control method, using a fuzzy rule library to adjust PID parameters, resulting in efficient parameter adjustment. A superconducting magnetic moment device is selected as the actuator, which generates a control torque through the interaction between the superconducting magnetic moment and the Earth's magnetic field for active suppression, providing flexible control.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and in particular to a nutation characteristic analysis method and an active suppression system for a spinning spacecraft. Background Art

[0002] With the huge increase in human demand for exploring and utilizing near-Earth space, traditional single-spin and dual-spin spacecraft are difficult to achieve established service goals such as continuous and high-speed in-orbit delivery of multiple payloads. Therefore, some spinning spacecraft came into being.

[0003] Some spinning spacecraft possess significant advantages, such as flexibility, controllability, and stability, enabling them to carry out more complex or specialized on-orbit missions, making them increasingly prominent in space resource utilization. Typical examples include NASA's proposal for space target transportation based on a space rotational momentum exchange tether and the Chinese Institute of Mechanics' proposal for low-orbit in-situ geomagnetic energy storage based on the principle of space-based rotational acceleration (Patents 201900773631.X and 201910774225.5). However, these partially spinning spacecraft often exhibit mass mutations, which can lead to nutation, posing significant challenges to the overall system's attitude stability.

[0004] Therefore, it is crucial to analyze the nutation characteristics of partially spinning spacecraft with mass mutation characteristics, and active nutation suppression methods need to be studied urgently. This work will lay a theoretical foundation for the development of special spacecraft dynamics and control. Summary of the Invention

[0005] The purpose of the present invention is to provide a method for analyzing the nutation characteristics of a spinning spacecraft and an active suppression system to solve the technical problem that the modeling method of the dynamic equations of partially spinning spacecraft with mass mutation characteristics in the prior art urgently needs to be studied.

[0006] In order to solve the above technical problems, the present invention specifically provides the following technical solutions:

[0007] A method for analyzing nutation characteristics of a spinning spacecraft comprises the following steps:

[0008] Step 100: Decompose the entire satellite system into a satellite platform body, a three-axis superconducting magnet system, a rotation separation device, a momentum exchange device, and a detachable payload according to its structural characteristics;

[0009] Step 200: Based on the orbit and structural characteristics of the satellite, a hypothesis for the derivation of the dynamic equations of the entire satellite is proposed to construct an ideal modeling environment;

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

[0011] The system center of mass orbit coordinate system OXYZ with the system center of mass as the 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 center of mass of the detachable load as the origin;

[0015] Step 400: Combined with the derivation assumptions of step 200, the system after load separation is regarded as a two-rigid body system. The angular momentum of the system is expressed as four parts, namely the angular momentum generated by the rotating separation device around its own center of mass, the angular momentum generated by the rotating separation device center of mass around the system center of mass, the angular momentum generated by the satellite platform body rotating around its own center of mass, and the angular momentum generated by the satellite platform body center of mass around the system center of mass, which are defined as And respectively calculate;

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

[0017] Transforming the resultant angular momentum based on the attitude conversion matrix between the reference system of the rotating separation device and the reference system of the satellite platform body to obtain the resultant angular momentum of the system relative to the center of mass;

[0018] Step 500: Calculate the system's relative angular momentum obtained in step 400 using the angular momentum theorem, and obtain the overall dynamic equation of the satellite system based on the derived assumptions.

[0019] Step 600: Set θ as the Euler rotation angle of the spacecraft from the inertial system to the local system, and use the ZYX rotation method to obtain the expression formula of θ:

[0020]

[0021] Then the Euler angle of the spacecraft attitude is calculated by integral solution, and the spacecraft attitude nutation angle expression formula is obtained through the attitude conversion matrix:

[0022] θ n =acos(cos(θ x )cos(θ z )+sin(θ x )sin(θ y )sin(θ z ));

[0023] Step 700 , based on the overall dynamic equation of the satellite system, a control variable is selected to simulate the shaping attitude motion to obtain the corresponding dynamic curves of the satellite under different variable conditions, and the different curves are fitted to obtain the variation of the maximum nutation angle of the satellite with each variable;

[0024] By calculating the spacecraft nutation amplitude and main nutation frequency when parameters change, the changes in nutation characteristics under different parameters can be analyzed.

[0025] As a preferred embodiment of the present invention,

[0026] In step 100;

[0027] The satellite platform body is composed of a rigid shell, internal rigid components and a momentum flywheel;

[0028] The momentum exchange device is composed of a stator and a rotor, and the stator is fixedly connected to the satellite platform body;

[0029] The three-axis superconducting magnet system is fixedly connected to the satellite platform body through the stator, and the three-axis superconducting magnet system can adjust the attitude of the satellite platform body through external torque;

[0030] The rotating separation device is fixedly mounted on the rotor, a slide rail for continuously conveying the separable load is provided on the rotating separation device, and a controllable instantaneous release mechanism is provided at the end of the slide rail;

[0031] The detachable load is a single rigid body slidably mounted on the slide rail, and the detachable load is launched through the controllable instantaneous release mechanism.

[0032] As a preferred solution of the present invention, the following assumptions are derived in step 200:

[0033] In the dynamic modeling and analysis, the satellite platform body, mass separation and rotation separation device, and superconducting magnet system are all assumed to be rigid bodies;

[0034] Ignore the mass of the momentum exchange device, the gap between the stator and rotor, the damping and other effects;

[0035] Assume that the mass separation rotating separation device and the load to be separated are separated instantaneously and there is no impact between them;

[0036] It is assumed that the satellite is in a stable state before separation.

[0037] As a preferred solution of the present invention, the specific steps of establishing multiple coordinate systems in step 300 are as follows:

[0038] System center-of-mass orbit coordinate system OXYZ: The origin O is at the center of mass of the system, the OZ axis points to the center of the earth, the OX axis is in the satellite orbit plane, perpendicular to the OZ axis and points in the direction of satellite motion, and the OY axis, OX axis and OZ axis form a right-handed coordinate system;

[0039] The satellite platform 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 directions of the axes are consistent with the directions of the corresponding axes of the orbital coordinate system OXYZ;

[0040] The rotating separation device body coordinate system OcXcYcZc: coordinate system origin O c Located at the intersection of the rotating axis of the rotating separation device and the rotating plane, it is assumed that at the initial moment, the directions of the axes of the rotating separation device and the coordinate system of the platform body are consistent;

[0041] The detachable load body coordinate system OdXdYdZd: coordinate system origin O d Located at the center of mass of the detachable load 5, the three coordinate axes before separation are parallel to the corresponding coordinate axes of the coordinate system of the rotating separation device body.

[0042] As a preferred solution of the present invention, in step 400, after the detachable load is separated, the angular momentum generated by the rotating separation device around its own center of mass is:

[0043]

[0044] in; The moment of inertia of the rotating separation device and the delivery load relative to the combined center of mass before separation;

[0045] ω is the relative angular velocity between the satellite platform body reference system and the center of mass orbit reference system;

[0046] Ω is the relative angular velocity between the optional device reference system and the satellite platform system.

[0047] As a preferred embodiment of the present invention, in step 400, the centroid of the rotating separation device is rotated around the system

[0048] The angular momentum generated by the combined center of mass is:

[0049]

[0050] Among them, m c is the mass of the rotating separation device;

[0051] r c is the radius vector of the center of mass of the rotating separation device relative to the total center of mass of the system;

[0052] ω is the relative angular velocity between the satellite platform body reference system and the center of mass orbit reference system;

[0053] Ω is the relative angular velocity between the optional device reference system and the satellite platform system.

[0054] As a preferred solution of the present invention, in step 400, the angular momentum generated by the satellite platform body rotating around its own center of mass is:

[0055]

[0056] in, The moment of inertia of the satellite platform body and the superconducting device relative to the combined center of mass;

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

[0058] As a preferred embodiment of the present invention, in step 400, the platform mass center is generated around the system mass center.

[0059] The resulting angular momentum:

[0060]

[0061] Among them, m b Combine the quality of the satellite platform body and superconducting device;

[0062] r b is the radius vector of the satellite platform's center of mass relative to the system's combined center of mass;

[0063] ω is the relative angular velocity between the satellite platform body reference system and the center of mass orbit reference system;

[0064] Ω is the relative angular velocity between the optional device reference system and the satellite platform system.

[0065] As a preferred solution of the present invention, in step 400, the time for calculating the moment of inertia of the center of mass of the rotating separation device relative to the total center of mass of the system is the moment when the separable load is separated from the rotating separation device, and it is assumed that there is no impact between the separable load and the rotating separation device.

[0066] As a preferred solution of the present invention, in step 700, the delivery mass m, the delivery arm length d, the distance e between the rotation plane of the rotating mechanism and the center of mass of the platform, and the relative angular velocity are selected as control variables to simulate the attitude motion of the entire satellite.

[0067] A system for actively suppressing nutation characteristics of a spinning spacecraft, using the above-mentioned method for analyzing nutation characteristics of a spinning spacecraft, comprises:

[0068] PID controller, fuzzy controller, superconducting magnetic moment, dynamic model and sensor;

[0069] Among them, the fuzzy controller inputs e = θ, ec = ω, and the outputs are the control torque pid parameters kp, ki, and kd. The inputs of e and ec are quantized and mapped to the digital interval of (-3, 3), and the triangular membership function is selected;

[0070] A fuzzy rule base is established for Δki, Δkp, and Δkd respectively, and the membership components of Δki, Δkp, and Δkd can be obtained. The fuzzy rule base Δki, Δkp, and Δkd are defuzzified by the PID controller to solve the size of each parameter and transmit it to the superconducting magnetic moment.

[0071] The superconducting magnetic moment inputs the control torque into the dynamic model, and the sensor feeds back the dynamic model monitoring information to the PID controller and the fuzzy controller, and so on, to achieve active suppression of the nutation characteristics of partial spinning spacecraft with mass mutation characteristics.

[0072] Compared with the prior art, the present invention has the following beneficial effects:

[0073] The present invention describes a method for analyzing the nutation characteristics of a partially spinning spacecraft, incorporating mass mutations. This analysis method can be used not only for on-orbit delivery spacecraft, but also for other spacecraft with mass mutation characteristics. The control method adopts a fuzzy PID control method, which uses a fuzzy rule library to adjust the PID parameters, resulting in high parameter adjustment efficiency. A superconducting magnetic moment device is selected as the actuator, which generates a control torque through the interaction between the superconducting magnetic moment and the geomagnetic field for active suppression, providing flexible regulation. BRIEF DESCRIPTION OF THE DRAWINGS

[0074] To more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for the embodiments or the description of the prior art. Obviously, the drawings described below are merely exemplary, and those skilled in the art can derive other implementation drawings based on the provided drawings without inventive effort.

[0075] Figure 1 A schematic diagram of a spinning spacecraft nutation characteristic analysis method and an active suppression system is provided for an embodiment of the present invention;

[0076] Figure 2 is the platform nutation characteristics under different delivery angular velocities;

[0077] Figure 3 The nutation characteristics of the platform under different delivery masses;

[0078] Figure 4 The nutation characteristics of the platform under different delivery arm lengths;

[0079] Figure 5 is the nutation characteristics of the platform under different eccentricities;

[0080] Figure 6 is the frequency domain characteristics of spacecraft attitude nutation;

[0081] Figure 7 System block diagram of the system for active suppression of nutation characteristics of spinning spacecraft;

[0082] Figure 8 is the functional relationship of the triangle membership function;

[0083] The numbers in the figure represent the following:

[0084] 1. Satellite platform body, 2. Three-axis superconducting magnet system;

[0085] 3 rotating separation device, 301 slide rail;

[0086] 4 momentum exchange device, 401 stator, 402 rotor;

[0087] 5. Separable load. DETAILED DESCRIPTION

[0088] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0089] like Figure 1 As shown, the present invention provides a method for analyzing the nutation characteristics of a spinning spacecraft, comprising the following steps:

[0090] Step 100: Decompose the entire satellite system into a satellite platform body, a three-axis superconducting magnet system, a rotation separation device, a momentum exchange device, and a detachable payload according to its structural characteristics.

[0091] Among them: the satellite platform body consists of a rigid shell, internal rigid components and a momentum flywheel.

[0092] The momentum exchange device consists of a stator and a rotor, and the stator is fixedly connected to the satellite platform body. The relative rotation direction and speed of the stator and rotor are controlled by the system.

[0093] The three-axis superconducting magnet system is fixedly connected to the satellite platform body through the stator. The three-axis superconducting magnet system can obtain kinetic energy by interacting with the geomagnetic field to inject energy into the system, and can also obtain external torque for adjusting the platform attitude or flywheel unloading.

[0094] The rotary separation device is fixedly installed on the rotor. A slide rail for continuously conveying the separable load is arranged on the rotary separation device, and a controllable instantaneous release mechanism is arranged at the end of the slide rail.

[0095] The detachable load is a single rigid body slidably mounted on a slide rail, and the detachable load is launched through a controllable instantaneous release mechanism.

[0096] Step 200: Based on the orbit and structural characteristics of the satellite, a hypothesis for the derivation of the dynamic equations of the entire satellite is proposed;

[0097] 1. In the dynamic modeling and analysis, the satellite platform body 1, the rotating separation device 3, and the three-axis superconducting magnet system 2 are all assumed to be rigid bodies.

[0098] 2. Ignore the mass of the momentum exchange device 4 and the gap between the stator 401 and the rotor 402, damping and other effects.

[0099] 3. Assume that the rotating separation device 3 and the separable load 5 are separated instantaneously without any impact between them.

[0100] 4. Assume that the satellite is in a stable state before separation.

[0101] Step 300: To facilitate the description of each component, the following coordinate system is introduced:

[0102] System center-of-mass orbit coordinate system OXYZ: The origin O is at the center of mass of the system, the OZ axis points to the center of the earth, the OX axis is in the satellite orbit plane, perpendicular to the OZ axis and points in the direction of satellite motion, and the OY axis, OX axis and OZ axis form a right-handed coordinate system;

[0103] Satellite platform body 1 coordinate system ObXbYbZb: The origin Ob is 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 directions of the axes are consistent with the directions of the corresponding axes of the orbital coordinate system OXYZ;

[0104] Coordinate system OcXcYcZc of the rotating separation device 3: The origin Oc of the coordinate system is located at the intersection of the rotation axis of the rotating separation device 3 and the rotation plane. It is assumed that at the initial moment, the directions of the axes of the rotating separation device 3 and the coordinate system of the platform body are consistent;

[0105] Coordinate system OdXdYdZd of the detachable load 5 body: the origin Od of the coordinate system is located at the center of mass of the detachable load 5, and the three coordinate axes before separation are parallel to the corresponding coordinate axes of the coordinate system of the rotating separation device 3 body.

[0106] In step 400, based on the assumptions derived in step 200, the system after load separation is regarded as a two-rigid body system (basic assumption 1 in step 200). The angular momentum of the system is expressed as four parts, namely, the angular momentum generated by the rotating separation device around its own center of mass, the angular momentum generated by the rotating separation device center of mass around the system 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 center of mass around the system center of mass, which are defined as And respectively calculate;

[0107] The resultant angular momentum is calculated for the angular momentum of the four systems mentioned above.

[0108] The following are the meanings of the symbols in the following formulas:

[0109] Symbol meaning

[0110] d Distance between the load and the center of mass of the rotating separation device

[0111] m c Rotating separation device mass

[0112] m b Satellite platform and superconducting device quality

[0113] Δm Mass of the load to be separated

[0114] h Distance between the rotation plane of the optional device and the center of mass of the platform

[0115] m1 Mass of the load to be separated

[0116] m2 Mass of load to be separated

[0117] e Mass of the load to be separated

[0118] u Mass of load to be separated

[0119] ω is the relative angular velocity between the satellite platform body reference frame and the center of mass orbit reference frame

[0120] Ω Relative angular velocity between the optional device reference system and the satellite platform system

[0121] r c Radius vector of the center of mass of the rotating separation device relative to the total center of mass of the system

[0122] r b Radius vector of the satellite platform's center of mass relative to the system's combined center of mass

[0123] Moment of inertia of the load and the rotating separation device relative to the combined center of mass before separation

[0124] The moment of inertia of the rotating separation device relative to its own center of mass

[0125] The moment of inertia of the satellite platform and superconducting device relative to the combined center of mass

[0126] The moment of inertia of the load relative to the center of mass of the optional device

[0127] The moment of inertia of the center of mass of the rotating separation device relative to the total center of mass of the system

[0128] C Attitude transformation matrix from the reference frame of the rotating separation device to the reference frame of the platform body

[0129] L Control torque generated by the interaction between superconducting magnetic moment and geomagnetic field

[0130] Among them, 1. The calculation process is as follows:

[0131] Assume that the inertia matrix of the delivered payload relative to its own center of mass is The distance from the center of mass c of the rotating separation device is d, then in the reference system of the rotating separation device, the moment of inertia of the delivery load relative to the center of mass c of the rotating separation device can be expressed as

[0132] in,

[0133]

[0134] For a rotating separation device, the center of mass offset due to load separation is:

[0135]

[0136] The inertia of the rotating separation device after mass separation can be expressed as:

[0137]

[0138] in, It is the moment of inertia of the rotating separation device and the delivery load relative to the combined center of mass before separation.

[0139] After the load is separated, the angular momentum generated by the rotating separation device around its own center of mass is:

[0140]

[0141] Similarly, the calculation results of the angular momentum generated by the center of mass of the rotating separation device around the combined center of mass of the system, the angular momentum generated by the platform rotating around its own center of mass, and the angular momentum generated by the center of mass of the platform around the combined center of mass of the system are as follows:

[0142] 2. Calculation

[0143] The angular momentum generated by the center of mass of the rotating separation device around the combined center of mass of the system:

[0144]

[0145] 3. Calculation

[0146] The angular momentum generated by the platform rotating around its own center of mass:

[0147]

[0148] 4. Calculation

[0149] The angular momentum generated by the center of mass of the platform around the center of mass of the system is:

[0150]

[0151] 5. Calculation and simplification of angular momentum

[0152] The resultant angular momentum can be expressed as:

[0153] in,

[0154]

[0155] is the moment of inertia of the center of mass of the rotating component relative to the combined center of mass at time t = 0;

[0156] C is the attitude transformation matrix between the reference system of the rotating separation device and the reference system of the platform body.

[0157] Since the load separation occurs instantaneously, there is no impact between the load and the rotating separation device (basic assumption 3 of step 200), and:

[0158]

[0159] Where: ΔH is the angular momentum loss due to load separation, expressed as:

[0160]

[0161] The initial kinetic energy of the system after load separation is:

[0162] Formula (10) E(t=0)=E0-ΔE;

[0163] where is the mechanical energy loss due to load separation.

[0164]

[0165] Step 500: Calculate the system's relative angular momentum obtained from formula (9) using the angular momentum theorem to obtain:

[0166]

[0167] Where: M is the sum of the interference torque and control torque acting on the satellite.

[0168] set up Simplifying, we can get:

[0169]

[0170] This equation is the overall dynamic equation of the satellite system.

[0171] The relative angular velocity between the satellite platform body 1 and the rotation separation device 3 is controlled by the momentum exchange device 4. According to the kinetic energy theorem, formula (13) can be obtained:

[0172]

[0173] Taking the derivative with respect to time, we get formula (14):

[0174]

[0175] The overall dynamic equation of the satellite system can be determined jointly by equations (11) and (14).

[0176] Step 600: Set θ as the Euler rotation angle of the spacecraft from the inertial system to the local system, and use the ZYX rotation method to obtain the expression formula of θ:

[0177]

[0178] Then the Euler angle of the spacecraft attitude is calculated by integral solution, and the spacecraft attitude nutation angle expression formula is obtained through the attitude conversion matrix:

[0179] θ n =acos(cos(θ x )cos(θ z )+sin(θ x )sin(θ y )sin(θ z ));

[0180] Step 700: Select control variables based on the overall dynamic equation of the satellite system to simulate the shaping attitude motion to obtain the corresponding dynamic curves of the satellite under different variable conditions, and fit the different curves to obtain the change of the maximum nutation angle of the satellite with each variable;

[0181] Based on the dynamic equation of the entire satellite system, the launch mass Δm, the launch arm length d, the distance e between the rotation plane of the rotating mechanism and the center of mass of the platform, and the relative angular velocity Ω are selected as control variables to simulate the attitude motion of the entire satellite, and the dynamic response curves of the satellite under different conditions are obtained: Figure 2 、 Figure 3 、 Figure 4 、 Figure 5 .

[0182] Furthermore, the fast Fourier transform (FFT) method is used to analyze the frequency domain characteristics of the spacecraft system nutation, and the main frequency of the spacecraft nutation is obtained. Figure 6 .

[0183] By calculating the spacecraft nutation amplitude and main nutation frequency when parameters change, the changes in nutation characteristics under different parameters can be analyzed.

[0184] A system for actively suppressing the nutation characteristics of a spinning spacecraft, using the above-mentioned method for analyzing the nutation characteristics of a spinning spacecraft, Figure 7 ,include:

[0185] PID controller, fuzzy controller, superconducting magnetic moment, dynamic model and sensor;

[0186] Among them, the fuzzy controller inputs e = θ, ec = ω, and the outputs are the control torque pid parameters kp, ki, and kd. The inputs of e and ec are quantized and mapped to the digital interval of (-3, 3), and the triangular membership function is selected;

[0187] A fuzzy rule base is established for Δki, Δkp, and Δkd respectively, and the membership components of Δki, Δkp, and Δkd can be obtained. The fuzzy rule base Δki, Δkp, and Δkd are defuzzified by the PID controller to solve the size of each parameter and transmit it to the superconducting magnetic moment.

[0188] The superconducting magnetic moment inputs the control torque into the dynamic model, and the sensor feeds back the dynamic model monitoring information to the PID controller and the fuzzy controller, and so on, to achieve active suppression of the nutation characteristics of partial spinning spacecraft with mass mutation characteristics.

[0189] Among them, the functional relationship of the triangle membership function refers to Figure 8 .

[0190] The fuzzy rule bases of Δki, Δkp, and Δkd are as follows:

[0191] ΔKp:

[0192]

[0193]

[0194] ΔKi:

[0195]

[0196] ΔKd:

[0197]

[0198] in,

[0199] The process of obtaining the triaxial superconducting magnetic moment control rate based on the size of each parameter is as follows:

[0200] Including: The torque m of the superconducting coil in the earth's magnetic field B is: L = m × B

[0201] It is known that the control torque expression in this system is:

[0202] L=kp·θ+ki·∫θdt+kd·ω;

[0203] get:

[0204] m=C·([B] * ·-L)=C·[[B] * ·(kp·θ+ki·∫θdt+kd·ω)];

[0205] in,

[0206]

[0207] C is the coordinate transformation matrix from the inertial system to this system.

[0208] This formula is the superconducting magnetic moment control rate. This equation is set in the control module of the controller, so that the nutation characteristics of some spinning spacecraft can be actively suppressed according to the input data and feedback data.

[0209] The above embodiments are merely exemplary embodiments of the present application and are not intended to limit the scope of the present application. The scope of protection of the present application is defined by the claims. Those skilled in the art may make various modifications or equivalent substitutions to the present application within the essence and scope of protection of the present application, and such modifications or equivalent substitutions shall also be deemed to fall within the scope of protection of the present application.

Claims

1. A method for analyzing nutation characteristics of a spinning spacecraft, characterized in that: The following steps are involved: Step 100: Decompose the entire satellite system into a satellite platform body (1), a three-axis superconducting magnet system (2), a rotation separation device (3), a momentum exchange device (4), and a detachable payload (5) according to its structural characteristics; Step 200: Based on the satellite's orbit and structural characteristics, a hypothesis is proposed for the derivation of the entire satellite's dynamic equations to construct an ideal modeling environment. The derivation hypothesis in step 200 is 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), the gap between the stator (401) and the rotor (402), and the damping effect; Assume that the rotating separation device (3) and the load to be separated are separated instantaneously without any impact between them; Assume that the satellite is in a stable state before separation; Step 300: To facilitate the description of each component, establish the following coordinate system: The system center of mass orbit coordinate system OXYZ with the system center of mass as the origin; O with the center of mass of the satellite platform body (1) as the origin b X b Y b Z b ; O with the center of mass of the rotating separation device (3) as the origin c X c Y c Z c ; O with the center of mass of the detachable load (5) as the origin d X d Y d Z d ; Step 400, combined with the derivation assumptions of step 200, the system after load separation is regarded as a two-rigid body system, and the system angular momentum is expressed as four parts, namely 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 total 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 total center of mass of the system, which are respectively defined as And respectively calculate; Calculate the resultant angular momentum of the four systems mentioned above; The angular momentum is transformed based on the attitude conversion matrix between the reference system of the rotation separation device (3) and the reference system of the satellite platform body (1) to obtain the angular momentum of the system relative to the center of mass; Step 500: Calculate the system's relative angular momentum obtained in step 400 using the angular momentum theorem, and obtain the overall dynamic equation of the satellite system based on the derived assumptions. Step 600: Set θ as the Euler rotation angle of the spacecraft from the inertial system to the local system, and use the ZYX rotation method to obtain the expression formula of θ: Then the Euler angle of the spacecraft attitude is calculated by integral solution, and the spacecraft attitude nutation angle expression formula is obtained through the attitude conversion matrix: i n =acos(cos(θ) x )cos(θ z )+sin(θ x )sin(θ y )sin(θ z )); Where ω is the relative angular velocity between the satellite platform body reference system and the center of mass orbit reference system; Step 700 , based on the overall dynamic equation of the satellite system, a control variable is selected to simulate the shaping attitude motion to obtain the corresponding dynamic curves of the satellite under different variable conditions, and the different curves are fitted to obtain the variation of the maximum nutation angle of the satellite with each variable; The spacecraft nutation amplitude and main nutation frequency are calculated when the parameters change, and the changes in nutation characteristics under different parameters are analyzed.

2. A method for analyzing nutation characteristics of a spinning spacecraft according to claim 1, characterized in that: In step 100; The satellite platform body (1) is composed of a rigid outer shell, internal rigid components and a momentum flywheel; The momentum exchange device (4) is composed of a stator (401) and a rotor (402), and the stator (401) is fixedly connected to the satellite platform body (1); The three-axis superconducting magnet system is fixedly connected to the satellite platform body (1) via the stator (401), and the three-axis superconducting magnet system controls the attitude of the satellite platform body (1) through an external torque; The rotating separation device (3) is fixedly mounted on the rotor (402), a slide rail (301) for continuously conveying the separable load (5) is provided on the rotating separation device (3), and a controllable instantaneous release mechanism is provided at the end of the slide rail (301); The detachable load (5) is a single rigid body slidably mounted on the slide rail (301), and the detachable load (5) is launched through the controllable instantaneous release mechanism.

3. The method for analyzing nutation characteristics of a spinning spacecraft according to claim 2, wherein: The specific steps of establishing multiple coordinate systems in step 300 are as follows: System center-of-mass orbit coordinate system OXYZ: The origin O is at the center of mass of the system, the OZ axis points to the center of the earth, the OX axis is in the satellite orbit plane, perpendicular to the OZ axis and points in the direction of satellite motion, and the OY axis, OX axis and OZ axis form a right-handed coordinate system; The satellite platform body (1) coordinate system O b X b Y b Z b : Origin O b Located at the mass center of the satellite platform body (1), when the three azimuth angles of the platform relative to the orbital coordinate system are zero, the directions of the axes are consistent with the directions of the corresponding axes of the orbital coordinate system OXYZ; The rotating separation device (3) has a body coordinate system O c X c Y c Z c : Coordinate system origin O c Located at the intersection of the rotation axis of the rotating separation device (3) and the rotating plane, and assuming that at the initial moment, the directions of the axes of the rotating separation device (3) and the coordinate axes of the platform body coordinate system are consistent; The detachable load (5) body coordinate system O d X d Y d Z d : Coordinate system origin O d Located at the centroid of the detachable load (5), the three coordinate axes before separation are respectively parallel to the corresponding coordinate axes of the main body coordinate system of the rotating separation device (3).

4. The method for analyzing nutation characteristics of a spinning spacecraft according to claim 3, wherein: In step 400, after the detachable load is separated (5), the angular momentum generated by the rotating separation device (3) around its own center of mass is: in; is the moment of inertia of the rotating separation device (3) and the delivery payload relative to the combined mass center before separation; Ω is the relative angular velocity between the optional device reference system and the satellite platform system.

5. The method for analyzing nutation characteristics of a spinning spacecraft according to claim 3, wherein: 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: Among them, m c is the mass of the rotating separation device; r c is the radius vector of the center of mass of the rotating separation device (3) relative to the total center of mass of the system; Ω is the relative angular velocity between the optional device reference system and the satellite platform body (1) system.

6. The method for analyzing nutation characteristics of a spinning spacecraft according to claim 3, wherein: 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 combined center of mass.

7. The method for analyzing nutation characteristics of a spinning spacecraft according to claim 3, wherein: In step 400, the angular momentum generated by the center of mass of the platform around the center of mass of the system is: Among them, m b The satellite platform body (1) and the superconducting device have a combined mass; r b is the radius vector of the mass center of the satellite platform (1) relative to the system mass center; Ω is the relative angular velocity between the optional device reference system and the satellite platform body (1) system.

8. The method for analyzing nutation characteristics of a spinning spacecraft according to claim 1, wherein: In step 400, the moment of inertia of the center of mass of the rotating separation device (3) relative to the total center of mass of the system is calculated at the moment when the detachable load (5) and the rotating separation device (3) are separated, and it is assumed that there is no impact between the detachable load (5) and the rotating separation device (3).

9. The method for analyzing nutation characteristics of a spinning spacecraft according to claim 1, wherein: In step 700, the delivery mass m, the delivery arm length d, the distance e between the rotation plane of the rotating mechanism and the center of mass of the platform, and the relative angular velocity are selected as control variables to simulate the attitude motion of the entire satellite.

10. A system for actively suppressing nutation characteristics of a spinning spacecraft, characterized in that: The method for analyzing nutation characteristics of a spinning spacecraft according to any one of claims 1 to 9 comprises: PID controller, fuzzy controller, superconducting magnetic moment, dynamic model and sensor; Among them, the fuzzy controller inputs e = θ, ec = ω, and the outputs are the control torque pid parameters kp, ki, and kd. The inputs of e and ec are quantized and mapped to the digital interval of (-3, 3), and the triangular membership function is selected; A fuzzy rule base is established for Δki, Δkp, and Δkd respectively, and the membership components of Δki, Δkp, and Δkd can be obtained. The fuzzy rule base Δki, Δkp, and Δkd are defuzzified by the PID controller to solve the size of each parameter and transmit it to the superconducting magnetic moment. The superconducting magnetic moment inputs the control torque into the dynamic model, and the sensor feeds back the dynamic model monitoring information to the PID controller and the fuzzy controller, and so on, to achieve active suppression of the nutation characteristics of partial spinning spacecraft with mass mutation characteristics.

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