Numerical simulation system and method for spacecraft formation and constellation operation control

Abstract

The present invention discloses a numerical simulation system and method for spacecraft formation and star cluster operation control. The system includes: a parameter setting module for setting required parameters and computing database data; a high-precision orbital dynamics module for realizing high-precision recursive calculation of spacecraft orbits and related data conversion; a navigation module for receiving absolute or relative positioning signals from a series of sensors and realizing filtering positioning; a guidance module for providing an ideal follow-up orbit state that changes with time to a control module; a control module for providing theoretical control output for spacecraft in formations and star clusters; a thruster module for simulating the thrust output of actual thrusters; and a parameter output module for setting the output of simulation results. The method can be used in the process of on-orbit flight missions of spacecraft formations and star clusters. The present invention can simulate the working process of navigation, guidance, control and actuators of multiple spacecraft in formations and star clusters, and has important practical value in the field of aerospace.

Description

Technical Field

[0001] The present invention belongs to the technical field of formation and cluster spacecraft orbit control simulation, and in particular relates to a numerical simulation system and method for spacecraft formation and cluster operation control. Background Art

[0002] With the rapid progress and development of modern aerospace technology, the concept and implementation of spacecraft miniaturization and modularization are becoming more mature. Micro-nano spacecraft with advantages such as light weight, low cost, short development cycle, high reliability and easy launch have been highly valued by many aerospace powers. A "virtual spacecraft" is constructed by multiple micro-spacecraft with the same or different functions flying in close formation, and through inter-satellite communication and task coordination, it can have the functions of a large spacecraft, and can even accomplish complex tasks that the latter cannot complete at this stage. Spacecraft formation and star cluster flight have broad application prospects in important scientific research fields such as scientific exploration, physical experiments, and electronic reconnaissance.

[0003] In order to quickly verify the overall reliability, accuracy and compatibility of the navigation, guidance, control and thrust subsystem functions of multiple spacecraft in formation and cluster flight missions, and ensure the smooth progress of the project in early planning, mid-term analysis and late verification, it is urgently needed to develop a software system and method that integrates navigation, guidance, control and thrust subsystems for unified numerical simulation. Summary of the invention

[0004] The object of the present invention is to provide a numerical simulation system and method for spacecraft formation and constellation operation control.

[0005] The technical solution for achieving the purpose of the present invention is: a numerical simulation system for spacecraft formation and constellation operation control, the numerical simulation system comprising a parameter setting module, a high-precision orbital dynamics module, a navigation module, a guidance module, a control module, a thruster module and a parameter output module;

[0006] The parameter setting module is used to set constants, variables and database data required by each module in the numerical simulation system;

[0007] The high-precision orbital dynamics module is used to realize high-precision orbital integral calculation of the spacecraft to obtain the absolute orbital position velocity state in the formation star cluster, and realize the conversion of multiple absolute and relative orbital data;

[0008] The navigation module is used to simulate the spacecraft in the formation and the constellation to receive absolute or relative positioning signals from a series of sensors and implement filtered positioning;

[0009] The guidance module is used to control the mission mode and provide the control modules of each spacecraft in the formation and the constellation with the ideal tracking orbit state that changes with time;

[0010] The control module is used to provide theoretical control outputs that meet constraint conditions for each spacecraft in the formation and the constellation based on navigation and guidance information;

[0011] The thruster module is used to simulate the thrust output of an actual thruster;

[0012] The parameter output module is used to set the output of simulation results.

[0013] Further, the parameter setting module includes a spacecraft basic parameter setting unit, a high-precision orbital dynamics module parameter setting unit, a navigation module parameter setting unit, a guidance module parameter setting unit, a control module parameter setting unit, a thruster module parameter setting unit and an output parameter requirement setting unit;

[0014] The spacecraft basic parameter setting unit is used to input the basic characteristic parameters of each spacecraft in the formation and the star cluster, including the number of spacecraft, the initial absolute orbit state of each spacecraft, the mass parameter of each spacecraft, the windward area of ​​each spacecraft, the drag coefficient of each spacecraft, the optical pressure area of ​​each spacecraft, and the thrust arrangement parameters of each spacecraft engine;

[0015] The high-precision orbital dynamics module parameter setting unit is used to input the parameters required for high-precision orbital dynamics, including high-order earth gravitational perturbation parameters, whether to consider the earth's polar axis drift nutation judgment flag, whether to consider the three-body perturbation judgment flag, whether to consider the solar light pressure judgment flag, whether to consider the atmospheric drag model judgment flag, and the initial mission simulation time;

[0016] The navigation module parameter setting unit is used to input navigation module calculation parameters, including absolute positioning and relative positioning selection flags, absolute positioning noise parameters, relative positioning noise parameters, and absolute or relative unscented Kalman filter parameters;

[0017] The guidance module parameter setting unit is used to input guidance module calculation parameters, including trajectory optimization algorithm parameters and ideal trajectory state;

[0018] The control module parameter setting unit is used to input control module calculation parameters, including main spacecraft absolute control parameters, model prediction control parameters, thruster default parameters, thruster amplitude constraint parameters and anti-collision constraint parameters;

[0019] The thruster module parameter setting unit is used to input the calculation parameters of each spacecraft thrust module, including thruster configuration parameters, thruster amplitude parameters, system delay parameters, thruster noise parameters, thruster fixed installation deviation, and spacecraft attitude disturbance parameters;

[0020] The output parameter requirement setting unit is used to set the simulation system output data parameters, including data time interval and output parameter type.

[0021] Furthermore, the high-precision orbital dynamics module includes a high-order earth gravitational acceleration calculation unit, an earth coordinate system transfer matrix calculation unit, a sun-moon three-body perturbation calculation unit, an atmospheric drag perturbation calculation unit, a thruster control force introduction unit, an integral operation unit, and a basic orbit parameter conversion unit;

[0022] The high-order earth gravity acceleration calculation unit is used to calculate the acceleration of the spacecraft in the earth orbit caused by the high-order earth non-spherical gravity;

[0023] The earth coordinate system transfer matrix calculation unit is used to calculate the transfer matrix from the earth inertial coordinate system to the geocentric coordinate system at the task simulation time point in the system;

[0024] The sun-moon three-body perturbation calculation unit is used to calculate the gravitational acceleration of the sun and the moon on the spacecraft in the earth's orbit;

[0025] The atmospheric resistance perturbation calculation unit is used to calculate the acceleration caused by the air resistance experienced by the spacecraft in the Earth's orbit;

[0026] The thruster control force introduction unit is used to introduce the acceleration of the spacecraft caused by the thrust provided by the thruster module of the simulation system;

[0027] The integral operation module is used to integrate and calculate the absolute position and velocity state at the next moment according to the real-time absolute position and velocity state of the spacecraft and various accelerations;

[0028] The basic orbit parameter conversion unit is used to convert the absolute orbit position velocity state of each spacecraft into the absolute orbit instantaneous element number, the absolute orbit average element number, the relative orbit position velocity state, and the relative orbit element number, etc.

[0029] Further, the navigation module includes an absolute navigation numerical simulation unit, a relative navigation numerical simulation unit, an absolute navigation filtering unit, a relative navigation filtering unit, and a track prediction unit;

[0030] The absolute navigation numerical simulation unit is used to simulate the measurement data information of the absolute position and speed of each spacecraft by the global navigation system or the global positioning system;

[0031] The relative navigation numerical simulation unit is used to simulate the measurement data information of relative position and speed of the relative optical camera, laser radar, and microwave radar distance and angle measurement information;

[0032] The absolute navigation filtering unit is used to filter and determine the absolute orbit position and velocity measurement data by using an unscented Kalman filtering algorithm;

[0033] The relative navigation filtering unit is used to filter and determine the relative orbit position and velocity measurement data by using an unscented Kalman filtering algorithm;

[0034] The orbit prediction unit is used to simulate the prediction of the satellite-borne short-term absolute orbit position velocity and relative orbit position velocity, and its output data will serve as the state prediction data source of the absolute navigation filtering module and the relative navigation filtering module.

[0035] Further, the guidance module includes a mission mode control unit and a nominal trajectory calculation unit;

[0036] The mission mode control unit is used to monitor, discipline and change the simulation system mission mode, and transmit the identification signal to the nominal trajectory calculation unit accordingly;

[0037] The nominal trajectory calculation unit is used to receive the nominal trajectory data and transmit it to the control module;

[0038] The control module includes a dynamic constraint calculation unit, a reference spacecraft absolute orbit control calculation unit and a secondary spacecraft relative orbit control calculation unit;

[0039] The dynamic constraint calculation unit is used to provide control calculation thruster defaults, thruster amplitude constraints and anti-collision constraints, and provide dynamic constraint parameters for the reference spacecraft absolute orbit control calculation unit and relative orbit control unit;

[0040] The reference spacecraft absolute orbit control calculation unit is used to calculate the optimal acceleration and thrust of the reference spacecraft under dynamic constraints and nominal trajectory guidance;

[0041] The secondary satellite spacecraft relative orbit control calculation unit is used to calculate the optimal acceleration and thrust of the secondary satellite spacecraft under dynamic constraints and nominal trajectory guidance.

[0042] Further, the thruster module includes a system delay simulation unit, a thruster noise fitting unit, a thruster amplitude constraint unit, a fixed assembly error setting unit, and a spacecraft attitude disturbance unit;

[0043] The system delay simulation unit is used to simulate the delay characteristics of each spacecraft navigation system, guidance system, control system, signal transmission and thruster response;

[0044] The thruster noise fitting unit is used to simulate the high-frequency noise and low-frequency noise of the thrust output by each spacecraft thruster;

[0045] The thruster amplitude constraint unit is used to constrain the amplitude of the thruster output;

[0046] The fixed assembly error setting unit is used to simulate the assembly angle of the thruster;

[0047] The spacecraft attitude disturbance unit is used to calculate the attitude disturbance of the spacecraft.

[0048] Furthermore, the parameter output module includes a parameter database unit and a parameter data image unit;

[0049] The parameter database unit is used to import the real-time data of the simulation process into the database at set time intervals;

[0050] The parameter data image unit is used to display the specific requirement data in the form of an image and save it in a working folder.

[0051] A numerical simulation method for spacecraft formation and constellation operation control based on the system comprises the following steps:

[0052] Step 1, setting all parameters required in the parameter setting module;

[0053] Step 2: After the system is running, the parameter settings are screened one by one. If the screening is wrong, the code setting interface will be re-entered and the error content will be informed. After the screening is passed, the next step will be entered;

[0054] Step 3, the parameter setting module determines the import database type according to the simulation parameter setting, imports the data required for simulation in the specified database, and establishes a calculation database for the system to perform calculations;

[0055] Step 4: The high-precision orbital dynamics module calculates the acceleration of the earth's gravity, the acceleration caused by air resistance, and the acceleration output by the thrusters of each spacecraft based on the absolute orbital position and velocity state data of the spacecraft, and performs dynamic integral calculations through the integral calculation unit to obtain the absolute orbital position and velocity state of the next step, and calculates the instantaneous root number, average root number, relative position and velocity state, and relative orbital root number of each spacecraft based on this;

[0056] Step 5, the navigation module determines whether to perform absolute navigation or relative navigation according to the navigation module parameters. When it is determined to be absolute navigation positioning, the process proceeds to step 6; when it is determined to be relative navigation positioning, the process proceeds to step 7;

[0057] Step 6, the absolute navigation numerical simulation unit of the navigation module applies the measurement noise of the characteristic parameters of the absolute sensor to the absolute orbit position velocity state, simulates the absolute orbit positioning measurement information, uses the orbit prediction unit to assist the unscented Kalman filter to determine the absolute position velocity state, and calculates the instantaneous root number, average root number, relative position velocity state, and relative orbit root number of each spacecraft after filtering, so as to provide input relative positioning data for the control system to realize high-precision relative orbit control;

[0058] Step 7, the relative navigation numerical simulation unit of the navigation module applies the measurement noise of the characteristic parameters of the relative sensor to the relative orbit position and velocity state, simulates the relative orbit positioning measurement information, uses the orbit prediction unit to assist the unscented Kalman filter to determine the relative position and velocity data, and calculates the relative orbit elements based on the relative position and velocity data to provide the input relative position and velocity data for the control system to realize high-precision relative orbit control;

[0059] Step 8: The mission mode control unit of the guidance module determines the mission status at this time and calculates whether it is necessary to change the mission or maintain the mission;

[0060] Step 9, the nominal trajectory calculation unit of the system guidance module extracts the nominal trajectory data from the corresponding task database and provides it to the control module after receiving the identification signal of the task mode control unit;

[0061] Step 10, the dynamic constraint calculation unit of the control module calculates the constraint conditions of the thruster default, thruster amplitude constraint and anti-collision constraint according to the parameter setting, and provides dynamic constraint parameters for the reference spacecraft absolute orbit control calculation unit and the relative orbit control unit;

[0062] Step 11, the reference spacecraft absolute orbit control calculation unit of the control module calculates the acceleration and thrust of the reference spacecraft under dynamic constraints and nominal trajectory guidance;

[0063] Step 12, the secondary satellite spacecraft relative orbit control calculation unit of the control module calculates the acceleration and thrust of the secondary satellite spacecraft under dynamic constraints and nominal trajectory guidance based on the model predictive control algorithm;

[0064] Step 13, the system delay simulation unit of the thruster module calculates and simulates the delay characteristics of the navigation system, guidance system, control system, signal transmission and thruster response of each spacecraft in real time, and feeds them back to the thruster output;

[0065] Step 14, the thruster noise fitting unit of the thruster module calculates and simulates the high-frequency noise and low-frequency noise of the thrust output of each spacecraft thruster in real time, and applies it to the thruster output;

[0066] Step 15, the thruster amplitude constraint unit of the thruster module constrains the amplitude of the thruster output so that the simulation output is not greater than the maximum output value of the thruster selected for the actual project;

[0067] Step 16, the fixed assembly error setting unit of the thruster module simulates the assembly angle of the thruster, calculates the thrust conversion matrix, and applies it to the thrust output value;

[0068] Step 17, the spacecraft attitude disturbance unit of the thruster module calculates the attitude disturbance parameters of the spacecraft, calculates the thrust conversion matrix, and applies it to the thrust output value;

[0069] Step 18, based on the thrust output value of step 17, the acceleration value of each spacecraft is calculated and transmitted to step 4 after coordinate transformation to realize the system closed loop.

[0070] Compared with the prior art, the present invention has the following significant advantages: 1) The system can simultaneously realize high-precision orbit recursive integral calculation of multiple spacecraft in formation and star cluster missions, including non-spherical perturbations of the earth, three-body perturbations, atmospheric drag, thrust acceleration of the thrust system, etc.; 2) The high-precision orbit dynamics module, navigation module, guidance module, control module and thruster module of the system can define the operating parameters of each module through the parameter setting module according to actual needs, so as to realize the research work of specific parameters of navigation, guidance and control, as well as the overall numerical simulation work of space missions, etc.; 3) The system has extremely high compatibility and can judge the number of formation spacecraft according to the initial parameters without changing the internal software framework, which is very suitable for the mission simulation of multiple spacecraft in formation and star cluster space missions; 4) The system contains a large number of basic calculation modules and can provide various types of simulation results Output, providing researchers with detailed process data so that they can conduct research on a certain parameter or module characteristic; 5) The system realizes the simulation of the entire closed-loop system including navigation, guidance, control and thrust subsystems in the spacecraft formation and constellation operation control to the greatest extent. The algorithms of each module adopt the current advanced calculation methods, such as the navigation module including the advanced unscented Kalman filter algorithm, the control module including the model predictive control, etc.; 6) The thruster module of the system fully considers the source of thrust error when the actual thruster is working, in order to judge whether the thruster performance can meet the mission requirements, to carry out thruster selection comparison, and to verify the matching of thruster characteristics to the navigation, guidance and control system. Provides an important simulation basis; 7) The system can construct the database required for operation according to the actual simulation parameters, thereby improving the efficiency and calculation speed of the overall simulation.

[0071] The present invention is further described in detail below in conjunction with the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0072] Figure 1This is the main structural framework diagram of the numerical simulation system for spacecraft formation and constellation operation control of the present invention.

[0073] Figure 2 This is the working principle and data flow diagram of the system parameter setting module of the present invention.

[0074] Figure 3 This is a flow chart of the recursive high-precision orbital dynamics model of multiple spacecraft and the basic orbital state calculation of the present invention.

[0075] Figure 4 It is a schematic diagram of the simulation process of the navigation module of the present invention.

[0076] Figure 5 It is a schematic diagram of the simulation process of the thruster module of the present invention. DETAILED DESCRIPTION

[0077] In order to make the purpose, technical solution and advantages of the present application more clearly understood, the present application is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present application and are not used to limit the present application.

[0078] In one embodiment, in combination Figure 1 , the present invention provides a numerical simulation system for spacecraft formation and constellation operation control, the numerical simulation system comprising a parameter setting module, a high-precision orbital dynamics module, a navigation module, a guidance module, a control module, a thruster module and a parameter output module;

[0079] The parameter setting module is used to set the constants, variables and database data required by each module in the numerical simulation system; Figure 2 ,The module includes a spacecraft basic parameter setting unit, a high-precision orbital dynamics module parameter setting unit, a navigation module parameter setting unit, a guidance module parameter setting unit, a control module parameter setting unit, a thruster module parameter setting unit and an output parameter requirement setting unit;

[0080] The spacecraft basic parameter setting unit is used to input the basic characteristic parameters of each spacecraft in the formation and the star cluster, including the number of spacecraft, the initial absolute orbit state of each spacecraft, the mass parameter of each spacecraft, the windward area of ​​each spacecraft, the drag coefficient of each spacecraft, the optical pressure area of ​​each spacecraft, and the thrust arrangement parameters of each spacecraft engine;

[0081] The high-precision orbital dynamics module parameter setting unit is used to input the parameters required for high-precision orbital dynamics, including high-order earth gravitational perturbation parameters, whether to consider the earth's polar axis drift nutation judgment flag, whether to consider the three-body perturbation judgment flag, whether to consider the solar light pressure judgment flag, whether to consider the atmospheric resistance model judgment flag, and the initial mission time;

[0082] The navigation module parameter setting unit is used to input navigation module calculation parameters, including absolute positioning and relative positioning selection flags, absolute positioning noise parameters, relative positioning noise parameters, and absolute or relative unscented Kalman filter parameters;

[0083] The guidance module parameter setting unit is used to input guidance module calculation parameters, including trajectory optimization algorithm parameters and ideal trajectory state;

[0084] The control module parameter setting unit is used to input control module calculation parameters, including main spacecraft absolute control parameters, model prediction control parameters, thruster default parameters, thruster amplitude constraint parameters and anti-collision constraint parameters;

[0085] The thruster module parameter setting unit is used to input the calculation parameters of each spacecraft thrust module, including thruster configuration parameters, thruster amplitude parameters, system delay parameters, thruster noise parameters, thruster fixed installation deviation, and spacecraft attitude disturbance parameters;

[0086] The output parameter requirement setting unit is used to set the simulation system output data parameters, including data time interval and output parameter type;

[0087] The high-precision orbital dynamics module is used to realize high-precision orbital integral calculation of spacecraft to obtain the absolute orbital position and velocity state in the formation star cluster, and realize the conversion of various absolute and relative orbital data; combined with Figure 3 ,This module includes a high-order earth gravitational acceleration calculation unit, an earth coordinate system transfer matrix calculation unit, a sun-moon three-body perturbation calculation unit, an atmospheric drag perturbation calculation unit, a thruster control force introduction unit, an integral operation unit, and a basic orbit parameter conversion unit;

[0088] The high-order earth gravity acceleration calculation unit is used to calculate the acceleration of the spacecraft in the earth orbit caused by the high-order earth non-spherical gravity;

[0089] The earth coordinate system transfer matrix calculation unit is used to calculate the transfer matrix from the earth inertial coordinate system to the geocentric coordinate system at the task simulation time point in the system;

[0090] The sun-moon three-body perturbation calculation unit is used to calculate the gravitational acceleration of the sun and the moon on the spacecraft in the earth's orbit;

[0091] The atmospheric resistance perturbation calculation unit is used to calculate the acceleration caused by the air resistance experienced by the spacecraft in the Earth's orbit;

[0092] The thruster control force introduction unit is used to introduce the acceleration of the spacecraft caused by the thrust provided by the thruster module of the simulation system;

[0093] The integral operation module is used to integrate and calculate the absolute position and velocity state at the next moment according to the real-time absolute position and velocity state of the spacecraft and various accelerations;

[0094] The basic orbit parameter conversion unit is used to convert the absolute orbit position velocity state of each spacecraft into the absolute orbit instantaneous element number, the absolute orbit average element number, the relative orbit position velocity state, and the relative orbit element number, etc.

[0095] The navigation module is used to simulate the formation and each spacecraft in the star cluster to receive the absolute or relative positioning signals of a series of sensors and realize filtering positioning; combined with Figure 4 , the module includes an absolute navigation numerical simulation unit, a relative navigation numerical simulation unit, an absolute navigation filtering unit, a relative navigation filtering unit, and a orbit prediction unit;

[0096] The absolute navigation numerical simulation unit is used to simulate the measurement data information of the absolute position and speed of each spacecraft by the global navigation system or the global positioning system;

[0097] The relative navigation numerical simulation unit is used to simulate the measurement data information of relative position and speed of the relative optical camera, laser radar, and microwave radar distance and angle measurement information;

[0098] The absolute navigation filtering unit is used to filter and determine the absolute orbit position and velocity measurement data by using an unscented Kalman filtering algorithm;

[0099] The relative navigation filtering unit is used to filter and determine the relative orbit position and velocity measurement data by using an unscented Kalman filtering algorithm;

[0100] The orbit prediction unit is used to simulate the prediction of the satellite-borne short-term absolute orbit position velocity and relative orbit position velocity, and its output data will serve as the state prediction data source of the absolute navigation filtering module and the relative navigation filtering module.

[0101] The guidance module is used to control the mission mode and provide the control modules of each spacecraft in the formation and the constellation with an ideal tracking orbit state that changes over time; it includes a mission mode control unit and a nominal trajectory calculation unit;

[0102] The mission mode control unit is used to monitor, discipline and change the simulation system mission mode, and transmit the identification signal to the nominal trajectory calculation unit accordingly;

[0103] The nominal trajectory calculation unit is used to receive the nominal trajectory data and transmit it to the control module.

[0104] The control module provides theoretical control outputs satisfying constraint conditions for each spacecraft in the formation and the star cluster based on navigation and guidance information; it includes a dynamic constraint calculation unit, a reference spacecraft absolute orbit control calculation unit and a secondary spacecraft relative orbit control calculation unit;

[0105] The dynamic constraint calculation unit is used to provide control calculation thruster defaults, thruster amplitude constraints and anti-collision constraints, and provide dynamic constraint parameters for the reference spacecraft absolute orbit control calculation unit and relative orbit control unit;

[0106] The reference spacecraft absolute orbit control calculation unit is used to calculate the optimal acceleration and thrust of the reference spacecraft under dynamic constraints and nominal trajectory guidance;

[0107] The secondary satellite spacecraft relative orbit control calculation unit is used to calculate the optimal acceleration and thrust of the secondary satellite spacecraft under dynamic constraints and nominal trajectory guidance.

[0108] The thruster module is used to simulate the thrust output of the actual thruster; combined with Figure 5 ,The module includes a system delay simulation unit, a thruster noise fitting unit, a thruster amplitude constraint unit, a fixed assembly error setting unit, and a spacecraft attitude disturbance unit;

[0109] The system delay simulation unit is used to simulate the delay characteristics of each spacecraft navigation system, guidance system, control system, signal transmission and thruster response;

[0110] The thruster noise fitting unit is used to simulate the high-frequency noise and low-frequency noise of the thrust output by each spacecraft thruster;

[0111] The thruster amplitude constraint unit is used to constrain the amplitude of the thruster output;

[0112] The fixed assembly error setting unit is used to simulate the assembly angle of the thruster;

[0113] The spacecraft attitude disturbance unit is used to calculate the attitude disturbance of the spacecraft.

[0114] The parameter output module is used to set the simulation result output; it includes a parameter database unit and a parameter data image unit;

[0115] The parameter database unit is used to import the real-time data of the simulation process into the database at set time intervals;

[0116] The parameter data image unit is used to display the specific requirement data in the form of an image and save it in a working folder.

[0117] In one embodiment, in combination Figures 2 to 5 , a numerical simulation method for spacecraft formation and constellation operation control based on the above system includes the following steps:

[0118] Step 1, setting all parameters required in the parameter setting module;

[0119] Step 2: After the system is running, the parameter settings are screened one by one. If the screening is wrong, the code setting interface will be re-entered and the error content will be informed. After the screening is passed, the next step will be entered;

[0120] Step 3, the parameter setting module determines the import database type according to the simulation parameter setting, imports the data required for simulation in the specified database, and establishes a calculation database for the system to perform calculations;

[0121] Step 4: The high-precision orbital dynamics module calculates the acceleration of the earth's gravity, the acceleration caused by air resistance, and the acceleration output by the thrusters of each spacecraft based on the absolute orbital position and velocity state data of the spacecraft, and performs dynamic integration through the integral operation unit to obtain the absolute orbital position and velocity state of the next step. At the same time, the instantaneous root number, average root number, relative position and velocity state, relative orbital root number, etc. of each spacecraft are calculated based on this;

[0122] Step 5, the navigation module determines whether to perform absolute navigation or relative navigation according to the navigation module parameters. When it is determined to be absolute navigation positioning, the process proceeds to step 6; when it is determined to be relative navigation positioning, the process proceeds to step 7;

[0123] Step 6, the absolute navigation numerical simulation unit of the navigation module applies the measurement noise of the characteristic parameters of the absolute sensor to the absolute orbit position velocity state, simulates the absolute orbit positioning measurement information, uses the orbit prediction unit to assist the unscented Kalman filter to determine the absolute position velocity state, and calculates the instantaneous root number, average root number, relative position velocity state, and relative orbit root number of each spacecraft after filtering, so as to provide input relative positioning data for the control system to realize high-precision relative orbit control;

[0124] Step 7, the relative navigation numerical simulation unit of the navigation module applies the measurement noise of the characteristic parameters of the relative sensor to the relative orbit position and velocity state, simulates the relative orbit positioning measurement information, uses the orbit prediction unit to assist the unscented Kalman filter to determine the relative position and velocity data, and calculates the relative orbit elements based on the relative position and velocity data to provide the input relative position and velocity data for the control system to realize high-precision relative orbit control;

[0125] Step 8: The mission mode control unit of the guidance module determines the mission status at this time and calculates whether it is necessary to change the mission or maintain the mission;

[0126] Step 9, the nominal trajectory calculation unit of the system guidance module extracts the nominal trajectory data from the corresponding task database and provides it to the control module after receiving the identification signal of the task mode control unit;

[0127] Step 10, the dynamic constraint calculation unit of the control module calculates the constraint conditions of the thruster default, thruster amplitude constraint and anti-collision constraint according to the parameter setting, and provides dynamic constraint parameters for the reference spacecraft absolute orbit control calculation unit and the relative orbit control unit;

[0128] Step 11, the reference spacecraft absolute orbit control calculation unit of the control module calculates the acceleration and thrust of the reference spacecraft under dynamic constraints and nominal trajectory guidance;

[0129] Step 12, the secondary satellite spacecraft relative orbit control calculation unit of the control module calculates the acceleration and thrust of the secondary satellite spacecraft under dynamic constraints and nominal trajectory guidance based on the model predictive control algorithm;

[0130] Step 13, the system delay simulation unit of the thruster module calculates and simulates the delay characteristics of the navigation system, guidance system, control system, signal transmission and thruster response of each spacecraft in real time, and feeds them back to the thruster output;

[0131] Step 14, the thruster noise fitting unit of the thruster module calculates and simulates the high-frequency noise and low-frequency noise of the thrust output of each spacecraft thruster in real time, and applies it to the thruster output;

[0132] Step 15, the thruster amplitude constraint unit of the thruster module constrains the amplitude of the thruster output so that the simulation output is not greater than the maximum output value of the thruster selected for the actual project;

[0133] Step 16, the fixed assembly error setting unit of the thruster module simulates the assembly angle of the thruster, calculates the thrust conversion matrix, and applies it to the thrust output value;

[0134] Step 17, the spacecraft attitude disturbance unit of the thruster module calculates the attitude disturbance parameters of the spacecraft, calculates the thrust conversion matrix, and applies it to the thrust output value;

[0135] Step 18, based on the thrust output value of step 17, the acceleration value of each spacecraft is calculated and transmitted to step 4 after coordinate transformation to realize the system closed loop.

[0136] Furthermore, in one embodiment, the filtering determination of the absolute / relative orbital position and velocity measurement data is implemented by an unscented Kalman filter algorithm, as follows:

[0137] First, for a given data [X, Y, U] T, where X = [x1,x2,x3,x4,x5,x6] T is the absolute / relative position velocity of the spacecraft, Y = [y1, y2, y3] T is the measured absolute / relative position obtained by the spacecraft, and U = [u1, u2, u3] is the absolute / relative acceleration value obtained by the spacecraft;

[0138] X k =f(X k-1 ,U k-1 )+q t ,q t ∈N(0,Q)

[0139] Y k =g(X k )+r t ,r t ∈N(0,R)

[0140] Where f(·) is the nonlinear dynamic transfer function of the state variable X under the action of U, k represents the time step, g(·) is the conversion function between the measured state Y and the state variable X, and q t and r t They are all white noises that obey Gaussian distribution. Q and R represent white noise q k and r k The variance of

[0141] Determine the value X by the absolute / relative position speed at time k-1 k-1 and the measured absolute / relative position value Y at time k k , to achieve the filtering determination value X of the absolute / relative position speed at time k k ; The specific calculation process is as follows:

[0142] (1) Weight allocation

[0143] The weights are allocated using the following formula:

[0144]

[0145]

[0146]

[0147] Where n is X k-1 The dimension of 2 (n+k)-n is a constant, α∈(0,1] and κ≥0 represent the scatter distribution range of the estimated value. The smaller α and κ are, the more concentrated the scatter points are, which can be used to adjust the impact of high-order terms on the model; β is an empirical parameter, represents the weight parameter of the i-th mean, represents the weight parameter of the i-th covariance, where i = 0, 1, 2, …, 2n;

[0148] (2) Calculate the estimated value

[0149] Get the determined value X of the state quantity at time k-1 k-1 Multiple estimates of is the i-th estimated value at time k-1, The number of is 2n+1, and the estimated calculation process is as follows:

[0150]

[0151]

[0152]

[0153]

[0154] In the formula, Cholesky(·) represents the Cholesky decomposition of the matrix, It is the covariance matrix of the previous moment, i.e., moment k-1. The initial value of this matrix is ​​given by the parameter setting, and the subsequent values ​​are given by iterative calculation; Represents a discrete matrix The i-th column array of , where i = 1, 2, ..., 2n; process (1) and process (2) become traceless transformation processes;

[0155] (3) Update the covariance value of the state quantity

[0156] The covariance value of the state quantity is updated using multiple estimated values ​​obtained by untraceable transformation. The calculation formula is as follows:

[0157]

[0158]

[0159] In the formula, is the estimated mean value at time k, is the covariance matrix of the estimated value at time k;

[0160] (4) Calculate the estimated value of the measured variable and update the covariance value of the measured variable

[0161] according to and Calculate the mean square error of the measurements and the measured estimated mean The calculation formula is as follows:

[0162]

[0163]

[0164]

[0165]

[0166]

[0167]

[0168] In the formula, is the i-th estimated value of the state variable at the current k moment, is the estimated value of the state The estimated mean of the measurements, is the mean square error of the measurements;

[0169] (4) Update the covariance values ​​of state variables and measurement variables

[0170] Estimated mean value based on measurements and the estimated mean values ​​of the state variables Calculate the covariance of the measured variables and the estimated values ​​of the state variables

[0171]

[0172] (4) Calculate the Kalman gain K and pass it through the real measurement variable Y k Get the determined state variable value X k :

[0173]

[0174]

[0175] (5) Update the covariance matrix of the state variables

[0176]

[0177] At this point, the unscented Kalman filter calculation method is completed, and the final output is the spacecraft absolute / relative position velocity state variable determination value X k , and the covariance matrix of the updated state variables

[0178] Furthermore, in one embodiment, the prediction of the orbit prediction unit is specifically as follows:

[0179] 1. The prediction process of the absolute orbit in the orbit prediction unit is as follows:

[0180] (1) Calculate all the acceleration values ​​a of the spacecraft all

[0181]

[0182] In the formula, a all Including the gravitational acceleration of the two bodies a TB , air resistance perturbation acceleration a AirDrag And the perturbation acceleration a of J2~J6 Ji , m is the mass of the spacecraft;

[0183] in,

[0184] The gravitational acceleration of the two bodies a TB :

[0185]

[0186] Atmospheric drag perturbation acceleration a AirDrag :

[0187]

[0188] J2 perturbation acceleration a J2 :

[0189]

[0190] J3 perturbation acceleration a J3 :

[0191]

[0192] J4 perturbation acceleration a J4 :

[0193]

[0194] J5 perturbation acceleration a J5 :

[0195]

[0196] J6 perturbation acceleration a J6 :

[0197]

[0198] In the formula, J i They represent the perturbation constant, R e =6378137m is the average radius of the earth, μ = 3.986×10 14 m 3 / s 2 is the Earth's gravitational constant, r is the distance between the spacecraft and the Earth's center, r I , r J , r KThey represent the X, Y, and Z axis positions of the spacecraft in the well-known J2000 coordinate system, ρ represents the atmospheric density at the location of the spacecraft, which is calculated from the database data, v represents the flight speed of the spacecraft relative to the earth, S is the frontal area of ​​the spacecraft, and C D represents the drag coefficient;

[0199] (2) Based on the acceleration values ​​calculated above, orbit prediction is performed using the RK7 (8) numerical integration method;

[0200] The acceleration value of the spacecraft obtained by this calculation method will be used to calculate the nonlinear dynamic transfer function f(·) of the unscented Kalman filter.

[0201] 2. The prediction process of the absolute orbit in the orbit prediction unit is as follows:

[0202] The relative orbit prediction is calculated based on the relative orbit elements, and the calculation formula is as follows:

[0203] δα k =Φ J2～J6 (α c ,dt)δα k-1 +Δδα k-1

[0204] In the formula, δα represents the relative orbital element number, Φ J2～J6 (α c ,dt) represents the time transfer matrix, Δδα k-1 It represents the change of relative orbital elements under the relative control input of control action U;

[0205] The detailed explanation of each parameter is as follows:

[0206] The relative orbital element δα is defined by the traditional Kepler orbital element as:

[0207] δα=f(α d ,a c ,i c )-f(α c ,a c ,i c )

[0208] =(δa,δλ,δe x ,δe y ,δi x ,δi y ) T

[0209] in:

[0210] f(α,a c ,i c )=(a / ac ,u+Ωcosi c ,e x ,e y ,,i,Ωsini)

[0211] Where α = (a, e, i, Ω, ω, M) is the traditional Kepler orbit element, a is the semi-major axis of the orbit, e is the orbit eccentricity, ω is the argument of perigee, M is the mean anomaly, u = ω + M is the argument of latitude, Ω is the right ascension of the ascending node, i is the orbit inclination, e is the orbital inclination, and x =ecosω,e y =esinω, subscript c represents the classical Kepler mean orbital element of the reference star, and subscript d represents the orbital element of the companion star;

[0212] The transfer matrix Φ of the relative orbital element δα with the time length dt under the non-spherical gravitational perturbation of J2~J6 J2～J6 (α c ,dt) is given by:

[0213]

[0214] φ 31 =-a(s ωf / c ω0 )G 4,a

[0215]

[0216]

[0217] φ 35 =-(s ωf / c ω0 )G 4,i

[0218] φ 41 =a(c ωf / c ω0 )G 4,a

[0219]

[0220]

[0221] φ 45 =(c ωf / c ω0 )G 4,i

[0222] In the formula, s (·) = sin(·), c (·) =cos(·), is the first-order derivative of the perigee argument ω under the influence of the perturbation of J2~J6, yes The partial derivative of the eccentricity e; where,

[0223]

[0224] Among them, q is the ordered set of perturbations to be considered g i is the absolute orbital element α of the primary star c The partial derivatives of an element i in are expressed as follows:

[0225]

[0226]

[0227]

[0228]

[0229] In the formula, and They are the right ascension of the ascending node Ω, the depression angle of the perigee ω, and the mean anomaly M, which is determined by the uniform zonal harmonics. The first-order derivative under the long-term perturbation caused by ;

[0230] The change of relative orbital elements under the action of U is shown in the following formula:

[0231]

[0232] In the formula, a c and n c They represent the semi-major axis and average orbital velocity of the reference star orbit, u=ω+M is the latitude argument of the main star, and Δδα represents the change value of the relative orbital element U=[u1,u2,u3] under the relative control input;

[0233] The calculation method is the nonlinear dynamic transfer function f(·) of the unscented Kalman filter in the relative orbit parameter determination.

[0234] The above shows and describes the basic principles, main features and advantages of the present invention. Those skilled in the art should understand that the present invention is not limited by the above embodiments, and the above embodiments and descriptions are only for explaining the principles of the present invention. Without departing from the spirit and scope of the present invention, any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims

1. A numerical simulation system for spacecraft formation and constellation operation control, characterized in that: The numerical simulation system includes a parameter setting module, a high-precision orbital dynamics module, a navigation module, a guidance module, a control module, a thruster module and a parameter output module; The parameter setting module is used to set constants, variables and database data required by each module in the numerical simulation system; The high-precision orbital dynamics module is used to realize high-precision orbital integral calculation of the spacecraft to obtain the absolute orbital position velocity state in the formation star cluster, and realize the conversion of multiple absolute and relative orbital data; The navigation module is used to simulate the spacecraft in the formation and the constellation to receive absolute or relative positioning signals from a series of sensors and implement filtered positioning; The guidance module is used to control the mission mode and provide the control modules of each spacecraft in the formation and the constellation with the ideal tracking orbit state that changes with time; The control module is used to provide theoretical control outputs that meet constraint conditions for each spacecraft in the formation and the constellation based on navigation and guidance information; The thruster module is used to simulate the thrust output of an actual thruster; The parameter output module is used to set the output of simulation results; The navigation module includes an absolute navigation numerical simulation unit, a relative navigation numerical simulation unit, an absolute navigation filtering unit, a relative navigation filtering unit, and a track prediction unit; The absolute navigation numerical simulation unit is used to simulate the measurement data information of the absolute position and speed of each spacecraft by the global navigation system or the global positioning system; The relative navigation numerical simulation unit is used to simulate the measurement data information of relative position and speed of the relative optical camera, laser radar, and microwave radar distance and angle measurement information; The absolute navigation filtering unit is used to filter and determine the absolute orbit position and velocity measurement data by using an unscented Kalman filtering algorithm; The relative navigation filtering unit is used to filter and determine the relative orbit position and velocity measurement data by using an unscented Kalman filtering algorithm; The orbit prediction unit is used to simulate the prediction of the short-term absolute orbit position velocity and relative orbit position velocity on board, and its output data will serve as the state estimation data source of the absolute navigation filter module and the relative navigation filter module; The absolute / relative orbit position and velocity measurement data are filtered and determined by the unscented Kalman filter algorithm as follows: First, for a given data [X, Y, U] T , where X = [x1,x2,x3,x4,x5,x6] T is the absolute / relative position velocity of the spacecraft, Y = [y1, y2, y3] T is the measured absolute / relative position obtained by the spacecraft, and U = [u1, u2, u3] is the absolute / relative acceleration value obtained by the spacecraft; X k =f(X k-1 ,U k-1 )+q t ,q t ∈N(0,Q) Y k =g(X k )+r t ,r t ∈N(0,R) Where f(·) is the nonlinear dynamic transfer function of the state variable X under the action of U, k represents the time step, g(g) is the conversion function between the measured state Y and the state variable X, and q t and r t They are all white noises that obey Gaussian distribution. Q and R represent white noise q k and r k The variance of Determine the value X by the absolute / relative position speed at time k-1 k-1 and the measured absolute / relative position value Y at time k k , to achieve the filtering determination value X of the absolute / relative position speed at time k k ; The specific calculation process is as follows: (1) Weight allocation The weights are allocated using the following formula: Where n is X k-1 The dimension of 2 (n+k)-n is a constant, α∈(0,1] and κ≥0 represent the scatter distribution range of the estimated value. The smaller α and κ are, the more concentrated the scatter points are, which can be used to adjust the impact of high-order terms on the model; β is an empirical parameter, represents the weight parameter of the i-th mean, represents the weight parameter of the i-th covariance, where i = 0, 1, 2, …, 2n; (2) Calculate the estimated value Get the determined value X of the state quantity at time k-1 k-1 Multiple estimates of is the i-th estimated value at time k-1, The number of is 2n+1, and the estimated calculation process is as follows: In the formula, Cholesky(·) represents the Cholesky decomposition of the matrix, It is the covariance matrix of the previous moment, i.e., moment k-1. The initial value of this matrix is ​​given by the parameter setting, and the subsequent values ​​are given by iterative calculation; Represents the estimate discrete point matrix The i-th column array of , where i = 1, 2, ..., 2n; process (1) and process (2) become traceless transformation; (3) Update the covariance value of the state quantity The covariance value of the state quantity is updated with the obtained multiple estimated values. The calculation formula is as follows: In the formula, is the estimated mean value at time k, is the covariance matrix of the estimated value at time k; (4) Calculate the estimated value of the measured variable and update the covariance value of the measured variable according to and Calculate the mean square error of the measurements and the measurement estimate mean The calculation formula is as follows: In the formula, is the i-th estimated value of the state variable at the current k-th moment, is the state estimate The estimated mean of the measurements, is the mean square error of the measurements; (5) Update the covariance values ​​of state variables and measurement variables Estimate the mean from the measurements and the estimated mean values ​​of the state variables Calculate the covariance of the measured variables and the estimated values ​​of the state variables (6) Calculate the Kalman gain K and pass it through the real measurement variable Y k Get the determined state variable value X k : (7) Update the covariance matrix of the state variables At this point, the unscented Kalman filter calculation method is completed, and the final output is the spacecraft absolute / relative position velocity state variable determination value X k , and the covariance matrix of the updated state variables 2. The numerical simulation system for spacecraft formation and constellation operation control according to claim 1, characterized in that: The parameter setting module includes a spacecraft basic parameter setting unit, a high-precision orbital dynamics module parameter setting unit, a navigation module parameter setting unit, a guidance module parameter setting unit, a control module parameter setting unit, a thruster module parameter setting unit and an output parameter requirement setting unit; The spacecraft basic parameter setting unit is used to input the basic characteristic parameters of each spacecraft in the formation and the star cluster, including the number of spacecraft, the initial absolute orbit state of each spacecraft, the mass parameter of each spacecraft, the windward area of ​​each spacecraft, the drag coefficient of each spacecraft, the optical pressure area of ​​each spacecraft, and the thrust arrangement parameters of each spacecraft engine; The high-precision orbital dynamics module parameter setting unit is used to input the parameters required for high-precision orbital dynamics, including high-order earth gravitational perturbation parameters, whether to consider the earth's polar axis drift nutation judgment flag, whether to consider the three-body perturbation judgment flag, whether to consider the solar light pressure judgment flag, whether to consider the atmospheric drag model judgment flag, and the initial mission simulation time; The navigation module parameter setting unit is used to input navigation module calculation parameters, including absolute positioning and relative positioning selection flags, absolute positioning noise parameters, relative positioning noise parameters, and absolute or relative unscented Kalman filter parameters; The guidance module parameter setting unit is used to input guidance module calculation parameters, including trajectory optimization algorithm parameters and ideal trajectory state; The control module parameter setting unit is used to input control module calculation parameters, including main spacecraft absolute control parameters, model prediction control parameters, thruster default parameters, thruster amplitude constraint parameters and anti-collision constraint parameters; The thruster module parameter setting unit is used to input calculation parameters of each spacecraft thrust module, including system delay parameters, thruster configuration parameters, thruster noise parameters, thruster amplitude parameters, thruster fixed installation deviation, and spacecraft attitude disturbance parameters; The output parameter requirement setting unit is used to set the simulation system output data parameters, including data time interval and output parameter type.

3. The numerical simulation system for spacecraft formation and constellation operation control according to claim 1, characterized in that: The high-precision orbital dynamics module includes a high-order earth gravitational acceleration calculation unit, an earth coordinate system transfer matrix calculation unit, a solar-moon three-body perturbation calculation unit, an atmospheric drag perturbation calculation unit, a thruster control force introduction unit, an integral calculation unit, and a basic orbital parameter conversion unit; The high-order earth gravity acceleration calculation unit is used to calculate the acceleration of the spacecraft in the earth orbit caused by the high-order earth non-spherical gravity; The earth coordinate system transfer matrix calculation unit is used to calculate the transfer matrix from the earth inertial coordinate system to the geocentric coordinate system at the task simulation time point in the system; The sun-moon three-body perturbation calculation unit is used to calculate the gravitational acceleration of the sun and the moon on the spacecraft in the earth's orbit; The atmospheric resistance perturbation calculation unit is used to calculate the acceleration caused by the air resistance experienced by the spacecraft in the Earth's orbit; The thruster control force introduction unit is used to introduce the acceleration of the spacecraft caused by the thrust provided by the thruster module of the simulation system; The integral calculation unit is used to integrate and calculate the absolute position and velocity state at the next moment according to the real-time absolute position and velocity state of the spacecraft and various accelerations; The basic orbit parameter conversion unit is used to convert the absolute orbit position velocity state of each spacecraft into the absolute orbit instantaneous element number, the absolute orbit average element number, the relative orbit position velocity state, and the relative orbit element number.

4. The numerical simulation system for spacecraft formation and constellation operation control according to claim 1, characterized in that: The orbit prediction unit's prediction is as follows:

1. The prediction process of the absolute orbit in the orbit prediction unit is as follows: (1) Calculate all the acceleration values ​​a of the spacecraft all In the formula, a all Including the gravitational acceleration of the two bodies a TB , air resistance perturbation acceleration a AirDrag And the perturbation acceleration a of J2~J6 Ji , m is the mass of the spacecraft; in, The gravitational acceleration of the two bodies a TB : Atmospheric drag perturbation acceleration a AirDrag : J2 perturbation acceleration a J2 : J3 perturbation acceleration a J3 : J4 perturbation acceleration a J4 : J5 perturbation acceleration a J5 : J6 perturbation acceleration a J6 : In the formula, J i They represent the perturbation constant, R e =6378137m is the average radius of the earth, μ = 3.986×10 14 m 3 / s 2 is the Earth's gravitational constant, r is the distance between the spacecraft and the Earth's center, r I , r J , r K They represent the X, Y, and Z axis positions of the spacecraft in the well-known J2000 coordinate system, ρ represents the atmospheric density at the location of the spacecraft, which is calculated from the database data, v represents the flight speed of the spacecraft relative to the earth, S is the frontal area of ​​the spacecraft, and C D represents the drag coefficient; (2) Based on the acceleration values ​​calculated above, orbit prediction is performed using the RK7 (8) numerical integration method; 2. The prediction process of the absolute orbit in the orbit prediction unit is as follows: The relative orbit prediction is calculated based on the relative orbit elements, and the calculation formula is as follows: yes k =Φ J2～J6 (a c ,dt)da k-1 +Dda k-1 In the formula, δα represents the relative orbital element number, Φ J2～J6 (α c ,dt) represents the time transfer matrix, Δδα k-1 Indicates the change of relative orbital elements under relative control input U; The detailed explanation of each parameter is as follows: The relative orbital element δα is defined by the traditional Kepler orbital element as: δα=f(α d ,a c ,i c )-f(a c ,a c ,i c ) =(δa,δλ,δe x ,the y ,di x ,di y ) T in: f(α,a c ,i c )=(a / a c ,u+Ωso c ,and x ,and y ,i,Ω) Where α = (a, e, i, Ω, ω, M) is the traditional Kepler orbit element, a is the semi-major axis of the orbit, e is the orbit eccentricity, ω is the argument of perigee, M is the mean anomaly, u = ω + M is the argument of latitude, Ω is the right ascension of the ascending node, i is the orbit inclination, e is the orbital inclination, and x =ecosω,e y =esinω, subscript c represents the classical Kepler mean orbital element of the reference star, and subscript d represents the orbital element of the companion star; The transfer matrix Φ of the relative orbital element δα with the time length dt under the non-spherical gravitational perturbation of J2~J6 J2～J6 (α c ,dt) is given by: f 31 =-a(s ωf / c ω0 )G 4,a f 35 =-(s ωf / c ω0 )G 4,i φ 41 =a(c ωf / c ω0 )G 4,a φ 45 =(c ωf / c ω0 )G 4,i In the formula, is the first-order derivative of the perigee argument ω under the influence of the perturbation of J2~J6, yes The partial derivative of the eccentricity e; where, Among them, q is the ordered set of perturbations to be considered g i is the absolute orbital element α of the primary star c The partial derivatives of an element i in are expressed as follows: Where U, U and U are the right ascension of the ascending node Ω and the depression angle of the perigee ω respectively, and the mean anomaly M is obtained by the uniform zonal harmonic The first-order derivative under the long-term perturbation caused by ; The change of relative orbital elements under the action of U is shown in the following formula: In the formula, a c and n c They represent the semi-major axis and average orbital velocity of the reference star orbit respectively, u=ω+M is the latitude argument of the main star, and Δδα represents the change value of the relative orbital elements U=[u1,u2,u3] under the relative control input.

5. The numerical simulation system for spacecraft formation and constellation operation control according to claim 1, characterized in that: The guidance module includes a mission mode control unit and a nominal trajectory calculation unit; The mission mode control unit is used to monitor, discipline and change the simulation system mission mode, and transmit the identification signal to the nominal trajectory calculation unit accordingly; The nominal trajectory calculation unit is used to receive the nominal trajectory data and transmit it to the control module; The control module includes a dynamic constraint calculation unit, a reference spacecraft absolute orbit control calculation unit and a secondary spacecraft relative orbit control calculation unit; The dynamic constraint calculation unit is used to provide control calculation thruster defaults, thruster amplitude constraints and anti-collision constraints, and provide dynamic constraint parameters for the reference spacecraft absolute orbit control calculation unit and relative orbit control unit; The reference spacecraft absolute orbit control calculation unit is used to calculate the optimal acceleration and thrust of the reference spacecraft under dynamic constraints and nominal trajectory guidance; The secondary satellite spacecraft relative orbit control calculation unit is used to calculate the optimal acceleration and thrust of the secondary satellite spacecraft under dynamic constraints and nominal trajectory guidance.

6. The numerical simulation system for spacecraft formation and constellation operation control according to claim 1, characterized in that: The thruster module includes a system delay simulation unit, a thruster noise fitting unit, a thruster amplitude constraint unit, a fixed assembly error setting unit, and a spacecraft attitude disturbance unit; The system delay simulation unit is used to simulate the delay characteristics of each spacecraft navigation system, guidance system, control system, signal transmission and thruster response; The thruster noise fitting unit is used to simulate the high-frequency noise and low-frequency noise of the thrust output by each spacecraft thruster; The thruster amplitude constraint unit is used to constrain the amplitude of the thruster output; The fixed assembly error setting unit is used to simulate the assembly angle of the thruster; The spacecraft attitude disturbance unit is used to calculate the attitude disturbance of the spacecraft.

7. The numerical simulation system for spacecraft formation and constellation operation control according to claim 1, characterized in that: The parameter output module includes a parameter database unit and a parameter data image unit; The parameter database unit is used to import the real-time data of the simulation process into the database at set time intervals; The parameter data image unit is used to display the specific requirement data in the form of an image and save it in a working folder.

8. A numerical simulation method for spacecraft formation and constellation operation control based on the system according to any one of claims 1 to 7, characterized in that: The following steps are involved: Step 1, setting all parameters required in the parameter setting module; Step 2: After the system is running, the parameter settings are screened one by one. If the screening is wrong, the code setting interface will be re-entered and the error content will be informed. After the screening is passed, the next step will be entered; Step 3, the parameter setting module determines the import database type according to the simulation parameter setting, imports the data required for simulation in the specified database, and establishes a calculation database for the system to perform calculations; Step 4: The high-precision orbital dynamics module calculates the acceleration of the earth's gravity, the acceleration caused by air resistance, and the acceleration output by the thrusters of each spacecraft based on the absolute orbital position and velocity state data of the spacecraft, and performs dynamic integration through the integral operation unit to obtain the absolute orbital position and velocity state of the next step, and calculates the instantaneous root number, average root number, relative position and velocity state, and relative orbital root number of each spacecraft based on this; Step 5, the navigation module determines whether to perform absolute navigation or relative navigation according to the navigation module parameters. When it is determined to be absolute navigation positioning, the process proceeds to step 6; when it is determined to be relative navigation positioning, the process proceeds to step 7; Step 6, the absolute navigation numerical simulation unit of the navigation module applies the measurement noise of the characteristic parameters of the absolute sensor to the absolute orbit position velocity state, simulates the absolute orbit positioning measurement information, uses the orbit prediction unit to assist the unscented Kalman filter to determine the absolute position velocity state, and calculates the instantaneous root number, average root number, relative position velocity state, and relative orbit root number of each spacecraft after filtering, so as to provide input relative positioning data for the control system to realize high-precision relative orbit control; Step 7, the relative navigation numerical simulation unit of the navigation module applies the measurement noise of the characteristic parameters of the relative sensor to the relative orbit position and velocity state, simulates the relative orbit positioning measurement information, uses the orbit prediction unit to assist the unscented Kalman filter to determine the relative position and velocity data, and calculates the relative orbit elements based on the relative position and velocity data to provide the input relative position and velocity data for the control system to realize high-precision relative orbit control; Step 8: The mission mode control unit of the guidance module determines the mission status at this time and calculates whether it is necessary to change the mission or maintain the mission; Step 9, the nominal trajectory calculation unit of the system guidance module extracts the nominal trajectory data from the corresponding task database and provides it to the control module after receiving the identification signal of the task mode control unit; Step 10, the dynamic constraint calculation unit of the control module calculates the constraint conditions of the thruster default, thruster amplitude constraint and anti-collision constraint according to the parameter setting, and provides dynamic constraint parameters for the reference spacecraft absolute orbit control calculation unit and the relative orbit control unit; Step 11, the reference spacecraft absolute orbit control calculation unit of the control module calculates the acceleration and thrust of the reference spacecraft under dynamic constraints and nominal trajectory guidance; Step 12, the secondary satellite spacecraft relative orbit control calculation unit of the control module calculates the acceleration and thrust of the secondary satellite spacecraft under dynamic constraints and nominal trajectory guidance based on the model predictive control algorithm; Step 13, the system delay simulation unit of the thruster module calculates and simulates the delay characteristics of the navigation system, guidance system, control system, signal transmission and thruster response of each spacecraft in real time, and feeds them back to the thruster output; Step 14, the thruster noise fitting unit of the thruster module calculates and simulates the high-frequency noise and low-frequency noise of the thrust output of each spacecraft thruster in real time, and applies it to the thruster output; Step 15, the thruster amplitude constraint unit of the thruster module constrains the amplitude of the thruster output so that the simulation output is not greater than the maximum output value of the thruster selected for the actual project; Step 16, the fixed assembly error setting unit of the thruster module simulates the assembly angle of the thruster, calculates the thrust conversion matrix, and applies it to the thrust output value; Step 17, the spacecraft attitude disturbance unit of the thruster module calculates the attitude disturbance parameters of the spacecraft, calculates the thrust conversion matrix, and applies it to the thrust output value; Step 18, based on the thrust output value of step 17, the acceleration value of each spacecraft is calculated and transmitted to step 4 after coordinate transformation to realize the system closed loop.

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