A method for constructing a dynamic library for rapid integration test of GNC software of high-orbit spacecraft

Abstract

The application provides a high-orbit spacecraft GNC software rapid integration test dynamics library construction method, comprising the following steps: (1) designing a function module of the spacecraft dynamics library; (2) programming to obtain the spacecraft dynamics library software; (3) determining the optimal configuration of an orbit dynamics module, an actuator module and a sensor module of the spacecraft dynamics library according to the function of the tested GNC software; (4) determining the initial configuration of an attitude dynamics module and a numerical solution module of the spacecraft dynamics library according to the function of the tested GNC software; (5) adopting the currently configured spacecraft dynamics library software and the tested GNC software to form a closed loop to constitute an integration test environment, and determining the optimal configuration of the attitude dynamics module of the spacecraft dynamics library through testing; and (6) adopting the currently configured spacecraft dynamics library software and the tested GNC software to form a closed loop to constitute an integration test environment, and determining the optimal configuration of the numerical solution module of the spacecraft dynamics library through testing. The method can optimize the calculation complexity of the spacecraft dynamics library, construct a spacecraft dynamics library with relatively small calculation complexity, significantly improve the integration test efficiency of the high-orbit spacecraft GNC software, is simple in engineering implementation, does not increase the cost of test equipment, and is particularly suitable for the integration test of the long-term function of the high-orbit spacecraft GNC software.

Description

Technical Field

[0001] This invention relates to a method for constructing a dynamics library for rapid integration testing of GNC software in high-orbit spacecraft, and particularly to a method for constructing a dynamics library for long-term functional integration testing of GNC software in high-orbit spacecraft, belonging to the field of spacecraft GNC software testing technology. Background Technology

[0002] According to the software development process, software integration testing is an essential step in ensuring the correctness of spacecraft GNC software implementation. The spacecraft integration testing environment consists of spacecraft dynamics library software and the GNC software under test. The spacecraft dynamics library software functions by simulating the spacecraft's attitude, orbit, and sensor / actuator functions using mathematical models of spacecraft attitude dynamics, orbital dynamics, and the functions of these components. This allows for the integration testing of the spacecraft GNC software on the ground.

[0003] However, extensive integration testing of spacecraft GNC software currently demonstrates that the spacecraft dynamics library software consumes an average of over 90% of the computation time per control cycle, while the onboard GNC software under test consumes less than 10% on average. Therefore, the computational complexity of the spacecraft dynamics library software is a major factor limiting the efficiency of software integration testing. This low efficiency and long processing time are particularly pronounced in the long-term functional integration testing of GNC software for high-orbit spacecraft.

[0004] Therefore, how to reduce the computation time of spacecraft dynamics library software and improve the integration and testing efficiency of high-orbit spacecraft GNC software, especially the long-term functionality of high-orbit spacecraft GNC software, is a pain point that urgently needs to be solved by those skilled in the art. Summary of the Invention

[0005] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a method for constructing a rapid integrated testing dynamics library for GNC software of high-orbit spacecraft.

[0006] The technical solution of this invention is: A method for constructing a rapid integrated test dynamics library for GNC software of high-orbit spacecraft includes the following steps: Step S1: Design the functional modules of the spacecraft dynamics library; Step S2: Program the spacecraft dynamics library software. Step S3: Determine the optimized configuration of the orbital dynamics module, actuator module, and sensor module of the spacecraft dynamics library based on the functions of the GNC software under test; Step S4: Determine the initial configuration of the attitude dynamics module and numerical solution module of the spacecraft dynamics library based on the functions of the GNC software under test; Step S5: The spacecraft dynamics library software with the current configuration and the GNC software under test are used to form an integrated test environment. The optimal configuration of the attitude dynamics module of the spacecraft dynamics library is determined through testing. Step S6: The spacecraft dynamics library software with the current configuration and the GNC software under test are used to form an integrated test environment. The optimal configuration of the numerical solution module of the spacecraft dynamics library is determined through testing.

[0007] Preferably, the functional modules of the spacecraft dynamics library in step S1 involve coordinate systems: J2000 geocentric inertial coordinate system (hereinafter referred to as inertial system), orbital coordinate system (hereinafter referred to as orbital system), spacecraft body coordinate system (hereinafter referred to as the body system), and Earth-fixed coordinate system (hereinafter referred to as Earth-fixed system). For the specific definitions of the four coordinate systems, please refer to reference [1] (Li Hengnian, Geostationary Satellite Orbit and Co-location Control Technology, National Defense Industry Press, 2010). The transformation matrix between coordinate systems is involved: the transformation matrix from the inertial system to the body system. Transformation matrix from orbital system to this system Transformation matrix from inertial frame to Earth-fixed frame Transformation matrix from inertial frame to orbital frame The calculation methods for the four transformation matrices can be found in reference [1] (Li Hengnian, Geostationary Satellite Orbit and Co-location Control Technology, National Defense Industry Press, 2010). The functional module design steps of the spacecraft dynamics library specifically include: 2.1 Orbital dynamics module design, specifically including: (1) Orbital time Orbital Time (s), where The initial time of the orbit (UTC, in seconds). Let be the current simulation time (s) for orbital dynamics, and require... No earlier than the base time start point Configure according to requirements; requires configuring the base time starting point. No earlier than 12:00:00 UTC on January 1, 2000, in seconds; (2) Orbital dynamics equations Under the influence of Earth's gravity and external perturbations, the orbital dynamics equations of the spacecraft in the inertial frame are as follows: in, and These represent position and velocity in the inertial frame, respectively. The distance from the Earth's center to the spacecraft. km 3 / s 2 The gravitational constant of Earth; Earth's gravitational perturbation (km / s) 2 ), Solar gravitational perturbation (km / s) 2 ), Lunar gravitational perturbation (km / s) 2 ), Solar radiation pressure perturbation (km / s) 2 ), For orbital jet perturbation (km / s) 2 ); The orbital dynamics module includes five perturbation terms, namely: Earth's gravitational perturbation Considering only the potential function of second-order Earth gravitational perturbation, i.e. ,in The geocentric longitude of the spacecraft. It is the geocentric latitude of the spacecraft. These are the position components of the spacecraft in the Earth-solid system, i.e. , km is the Earth's average equatorial radius, and the coefficient of the second-order zonal harmonic term. Second-order harmonic coefficient Geographic longitude of the second-order field harmonic term principal axis deg; Solar gravitational perturbation ,in km 3 / s 2 The gravitational constant of the Sun, This represents the vector (km) from the spacecraft to the sun with a magnitude of ; Represents the vector (km) from Earth to the Sun with a magnitude of ,in Given from the ephemeris; Lunar gravitational perturbation ,in km 3 / s 2 The gravitational constant of the moon, Represents the vector (km) from the spacecraft to the moon with a magnitude of ; Represents the vector (km) from Earth to the Moon with a magnitude of ,in Given from the ephemeris; Solar radiation pressure perturbation ,in For sun-receiving factors, Indicates the area of ​​the ground shadow. Indicates the area under illumination. The surface reflectance of the spacecraft is , and the solar radiation pressure intensity is . (N / m) 2 ), Solar radiation pressure surface mass ratio (m) 2 / kg), bound according to spacecraft configuration and material properties; Orbit-controlled jet perturbation The thrust vector of the orbital control engine in the body is Unit N, modulus is , The real-time mass (kg) of the entire satellite, and , , , This is the initial mass of the entire star. Indicates the rail control engine in the first The jet pulse width (s) for each control cycle is determined by the GNC software under test; and Indicates the first Specific impulse of individual attitude control thrusters This indicates the specific impulse of the rail control engine. Represents the acceleration due to gravity at the Earth's surface (m / s²). 2 ); (3) Orbital dynamics module configuration parameters, including: control period (Require (Unit: seconds); initial orbital time (UTC) (Configuration required no earlier than) (Unit: seconds); reference time start point (Configuration must be no earlier than 12:00:00 UTC on January 1, 2000), unit: seconds; reference time offset. Indicates the starting point of the reference time. The number of seconds relative to 12:00:00 UTC on January 1, 2000 (configuration required). ), unit s; sun exposure factor Take in the shadow area Take in the lighted area Spacecraft surface reflectivity Typically, 1.5 is used; solar radiation pressure surface mass ratio Unit m 2 / kg; initial mass of the spacecraft Unit: kg; Thrust vector of the orbital control engine in this system. Unit: N; Specific impulse of the rail control engine Unit: s; Initial position of the inertial frame. Unit: km, Initial velocity of the inertial frame Unit: km / s; Track dynamics module storage period ( (positive integer), unit s; allowable error limit of inertial frame position. Unit: km; Permissible error limit of inertial frame velocity The unit is km / s; (4) Variables recorded by the orbital dynamics module, including: the number of orbital times. The unit is s, where The number of positions in the inertial frame. Unit: km; Number of inertial frame velocities The unit is km / s; 2.2 Attitude dynamics module design, specifically including: (1) Posture kinematic equations ( , and Values (But they are all different) The kinematic equations of the spacecraft attitude under the transition sequence are: in, These represent the roll angle, pitch angle, and yaw angle of the spacecraft relative to its orbital system, respectively. This represents the component of the spacecraft's instantaneous rotational speed relative to the orbital system within this system. Represents Euler angular rate and Relationship matrix, 6 different transformations The expression for is found in Chapter 2 of reference [2] (Tu Shancheng, ed. Satellite Attitude Dynamics and Control. Aerospace Press. 2001); , The rotational velocity of the body relative to the inertial frame is expressed in this system. The rotational speed of the orbital frame relative to the inertial frame is expressed in the orbital frame, where the Earth's gravitational constant is... km 3 / s 2 , The semi-major axis of the spacecraft's orbit (km). This represents the transformation matrix from the orbital system to the home system, while Indicate Euler angles at The attitude matrix corresponding to the transposed order. For the specific expression, please refer to Chapter 2 of reference [2] (Tu Shancheng, ed. Satellite Attitude Dynamics and Control. Aerospace Press. 2001).

[0008] (2) Attitude dynamics equations Flexible spacecraft dynamics (with symmetrical solar panels, considering only the effect of panel vibration on the rotation of the central rigid body, ignoring the effect of panel vibration on the translation and rotation of the central rigid body, and ignoring the effect of panel rotation on the rotation of the central rigid body) has two configurable types, as follows: (i) Momentum wheel control (referred to as wheel-controlled attitude control) attitude dynamics of flexible spacecraft in, For the entire star's inertia matrix (relative to the center of mass); The installation matrix of the momentum wheel assembly. ; , The moment of inertia matrix of the momentum wheel assembly, in kg. m 2 ,in The number of momentum wheels, where For the first The moment of inertia of a momentum wheel; Let be the speed vector of the momentum wheel assembly, in rad / s, where For the first The rotational speed of the momentum wheel; and This represents the flexible coupling coefficient matrix of the vibrations of the left and right solar panels on the rotation of the spacecraft (central rigid body). ,in The modal order; and These are the modal coordinate vectors for the left and right solar panels; and This is the diagonal matrix of the modal frequencies of the left and right solar panels; and This is a diagonal matrix of the modal damping coefficients for the left and right solar panels, with the diagonal elements typically set to 0.005. Indicates the wheel control attitude control torque; The driving torque vector of the momentum wheel assembly (calculated by the GNC software under test), in Nm; The driving torque noise of the momentum wheel assembly, where This indicates that the mean is 0 and the variance is... Noise level, in Nm; This is the diagonal matrix in the direction of the momentum wheel assembly's rotational speed. ; This is the set value of the dynamic friction torque of the momentum wheel assembly, in Nm; The damping friction coefficient of the momentum wheel is given in Nm / rpm; the nominal speed vector of the momentum wheel assembly is given. (Configurable), unit rpm; the operating speed vector range of the momentum wheel assembly is... (Configurable), where , The unit is rpm; Represents the external disturbance torque, where The solar pressure torque is expressed in triaxial form within this system, in Nm. ,and , and , This is the optical pressure torque coefficient, which is related to the geometry and mass distribution characteristics of the spacecraft. The value is the representation of the track control interference torque in this system, and is a bound value in Nm; (ii) Jet control (or jet attitude control) attitude dynamics of flexible spacecraft Among them, the The lever arm vector of the attitude control thruster The thrust vectors of the three axes in this system , Indicates the first The jet pulse sequence of an attitude control thruster is defined as follows: The assignment logic of the attitude control thruster command in the tested GNC software is determined. Indicates the first The attitude control thruster in the first Jet pulse width per control cycle, in seconds. ; (3) Attitude dynamics module configuration parameters, including: sequence transition It includes 6 types of sequence transformations: , , , , , Inertia matrix of the entire star (Relative center of mass), unit kg.m 2 , Flexible coupling coefficient matrix of left and right solar panel vibrations to the rotation of the central rigid body and , Modal order Modal frequencies of the left and right solar panels and Diagonal matrix of modal damping coefficients for left and right solar panels and The diagonal elements are generally taken as 0.005; the number of momentum wheels Moment of inertia of the momentum wheel , Unit: kg m 2; Moment of inertia matrix of momentum wheel assembly , Unit: kg m 2 Installation matrix of momentum wheel assembly dimensionality Momentum wheel assembly driving torque noise Unit: Nm; Momentum wheel assembly speed measurement noise Unit: rpm; dynamic friction torque of momentum wheel assembly (set value) The unit is Nm; the damping friction coefficient of the momentum wheel. Unit: Nm / rpm; Number of attitude control thrusters Attitude control thruster nozzle installation position , , unit m; attitude control thruster nozzle vector angle (angle with the positive directions of the three axes of the body). , Unit: deg; Nominal force of attitude control thruster , Unit N; the first The specific impulse of the individual attitude control thruster , Unit: seconds; Initial attitude angle , unit: rad; initial attitude angular velocity Unit: rad / s; Attitude dynamics module data storage cycle ( (positive integer), unit s; attitude angle tolerance limit Unit: rad; allowable error limit for attitude angular velocity. The unit is rad / s; (4) Variables recorded by the attitude dynamics module, including: the stored values ​​of the three-axis attitude angles. ,in , in rad, where ; Storage of three-axis attitude angular velocities Unit: rad / s; Total number of optimization tests for the attitude dynamics module. Dimensionless; 2.3 Actuator module design, specifically including: (1) Momentum wheel design: based on Calculate The rotational speed of the momentum wheel assembly composed of momentum wheels (rad / s), then we get Measuring the rotational speed of the momentum wheel (rpm), output to the GNC software under test; where For momentum wheel assembly speed measurement noise, The mean is 0 and the variance is Noise (rpm) ; (2) Attitude control thruster design: according to Calculate the attitude control torque (Nm) acting on the three axes of the spacecraft's main system during the current control cycle, and then drive the spacecraft's attitude evolution; (3) Track control engine design: according to Calculate the orbital control force (N) applied to the three axes of the spacecraft's inertial frame during the current control cycle, and then drive the spacecraft's orbital evolution; 2.4 Sensor module design, specifically including: (1) Star sensor design: For star sensors The attitude dynamics module outputs three-axis attitude angles. Unit: rad, with added three-axis attitude equivalent measurement noise. The unit is rad. ,according to The attitude matrix of this system relative to the orbital system is obtained by transposing the order. Then according to Calculation of star sensitivity Measured spacecraft attitude matrix Convert to the corresponding quaternion Then we get Xingmin Measured spacecraft attitude quaternions ;in , and The mean is 0 and the variance is noise, Xingmin Installation matrix (configurable), number of star sensors (Configurable) ; (2) Gyroscope design: according to Calculate the first The angle increment of the gyroscope only , in rad; where The three-axis attitude angular velocity output by the attitude dynamics module, in rad / s; For the first Add a constant drift value to the gyroscope, in rad / s; With a mean of 0 and a variance of The random drift, in rad / s; Indicates the first Mounting vector for only the gyroscope input axis. To control the period, the number of gyroscopes (Configurable), and the configuration must ensure that at least 3 gyroscopes are not coplanar. ; (3) Sensor module configuration parameters, including: number of star sensors Starsense Installation Matrix , Three-axis attitude equivalent measurement noise , Number of spinning tops ; Gyroscope mounting vector ,in ; Gyroscope constant drift ,in ; Gyroscope measurement noise ,in ; 2.5 Numerical solution module design, specifically including: (1) Numerical solution algorithm, using the fourth-order Runge-Kutta algorithm to realize the numerical solution of the spacecraft attitude dynamics equation and orbital dynamics equation; (2) Numerical solution step size The unit is ms, which is the configuration parameter to be optimized; (3) Configuration parameters of the numerical solution module, including: numerical solution step size Unit: milliseconds; Dynamics library optimization test time. That is, optimizing the time interval of the dynamics library through testing. ,in Based on the configuration of the tested GNC software, the unit is seconds; the actual duration of the tested GNC software function. That is, the actual time interval of the tested GNC software function is ,in According to the functional configuration of the tested GNC software, the unit is s; (4) Variables recorded by the numerical solution module, including: the total number of optimization tests of the numerical solution module. Dimensionless; 2.6 Interface design between functional modules of the spacecraft dynamics library, including: (1) Input of the orbital dynamics module, including: the current state of the orbital dynamics equation (output of the numerical solution module), and the orbital control force (output of the actuator module); (2) Output of the orbital dynamics module, including: the state of the orbital dynamics equation of the previous long time step (for the numerical solution module); (3) Input of the attitude dynamics module, including: the current state of the attitude dynamics equation (output of the numerical solution module), and the attitude control torque (output of the actuator module); (4) Output of the attitude dynamics module, including: the state of the attitude dynamics equation of the previous long time step (for the numerical solution module), and the spacecraft attitude angle and attitude angular velocity. (5) Outputs of the actuator module, including: attitude control torque (to attitude dynamics module), orbit control force (to orbit dynamics module); (6) Inputs of the sensor module, including: spacecraft attitude angle and attitude angular velocity (outputs of attitude dynamics module); (7) Inputs of the numerical solution module, including: attitude dynamics equation state of the previous long time step (output of attitude dynamics module), orbit dynamics equation state of the previous long time step (output of orbit dynamics module); (8) Outputs of the numerical solution module, including: attitude dynamics equation state of the current time step (to attitude dynamics module), orbit dynamics equation state of the current time step (to orbit dynamics module).

[0009] Preferably, the spacecraft dynamics library software obtained by programming in step S2 refers to the following: the five functional modules designed according to steps 2.1 to 2.5, and the interfaces between the five functional modules of the dynamics library designed according to step 2.6, are implemented using a programming language consistent with the programming language of the GNC software under test; the interaction interface between the spacecraft dynamics library software and the GNC software under test includes: (1) the orbital time output by the orbital dynamics module. Spacecraft position and speed (GNC software under test); (2) Quaternion of spacecraft attitude measured by star sensor output by sensor module Angular increment measured by gyroscope (GNC software under test); (3) The momentum wheel output by the actuator module measures the rotational speed. (GNC software under test); (4) Wheel control attitude control torque input to the actuator module Attitude control jet pulse width Orbit control jet pulse width (Output from the tested GNC software).

[0010] Preferably, the optimized configuration of the orbital dynamics module, actuator module, and sensor module of the spacecraft dynamics library in step S3 is as follows: (1) Optimized configuration of orbital dynamics module: The orbital dynamics module is configured with Earth's central gravitational acceleration. If the tested GNC software function involves orbit control jet thrust, then the orbit dynamics module should be configured with the orbit control jet thrust item. Otherwise, do not configure. If the tested GNC software function involves longitude... Peaceful longitude change rate Then the orbital dynamics module is configured with Earth's gravitational perturbation terms. Otherwise, do not configure. If the GNC software function being tested involves tilt angle... Right ascension of ascending node Then the orbital dynamics module is configured with solar gravitational perturbation terms. and lunar gravitational perturbation terms Otherwise, do not configure. and If the tested GNC software function involves eccentricity... Perigeal argument Then the orbital dynamics module is configured with solar radiation pressure perturbation terms. Otherwise, do not configure. ; (2) Optimized configuration of actuator module: Determine the configuration parameters of the actuator module based on the functions of the GNC software under test: number of momentum wheels Moment of inertia of the momentum wheel , Unit: kg m 2 ; Moment of inertia matrix of momentum wheel assembly , Unit: kg m 2 Installation matrix of momentum wheel assembly dimensionality Momentum wheel assembly driving torque noise Unit: Nm; Momentum wheel assembly speed measurement noise Unit: rpm; dynamic friction torque of momentum wheel assembly (set value) The unit is Nm; the damping friction coefficient of the momentum wheel. Unit: Nm / rpm; Number of attitude control thrusters Attitude control thruster nozzle installation position , , unit m; attitude control thruster nozzle vector angle (angle with the positive directions of the three axes of the body). , Unit: deg; Nominal force of attitude control thruster , Unit N; the first The specific impulse of the individual attitude control thruster , Unit: s; Thrust vector of the orbital control engine in this system. Unit: N; Specific impulse of the rail control engine Unit: s; (3) Sensor module optimization configuration: Determine the sensor module configuration parameters based on the functions of the GNC software under test: number of star sensors Starsense Installation Matrix , Three-axis attitude equivalent measurement noise , Number of spinning tops ; Gyroscope mounting vector ,in ; Gyroscope constant drift ,in ; Gyroscope measurement noise ,in .

[0011] Preferably, the initial configuration steps for the attitude dynamics module and numerical solution module of the spacecraft dynamics library in step S4 are as follows: (1) Initial configuration of attitude dynamics module: If the GNC software under test involves spacecraft position and velocity, but not spacecraft attitude angle and attitude angular velocity, and wheel control is used for attitude control, then the attitude dynamics module is configured directly. rad, rad / s, meaning the attitude dynamics module no longer performs numerical calculations; if the tested GNC software function involves spacecraft attitude angles and attitude angular velocities, and wheel-controlled attitude control is used without orbit control, then rigid body spacecraft attitude dynamics (i.e., configuration) is configured. and ), wheel-controlled attitude control torque and solar pressure torque; if the measured GNC software function involves spacecraft attitude angles and attitude angular velocities, and wheel-controlled attitude control and jet orbital control are used, then rigid body spacecraft attitude dynamics should be configured (i.e., configuration of rigid body spacecraft attitude dynamics). and ), wheel control attitude control torque, solar pressure torque, and orbital control disturbance torque; if the tested GNC software function uses jet attitude control and has no orbital control, then configure the flexible spacecraft attitude dynamics and jet attitude control torque, including configuring the modal order. External disturbance torque If the tested GNC software function uses jet attitude control and jet orbit control, then configure the flexible spacecraft attitude dynamics, jet attitude control torque, and orbit control disturbance torque, including configuring the modal order. External disturbance torque ; (2) Initial configuration of numerical solution module: The numerical solution algorithm adopts the fourth-order Runge-Kutta algorithm, and the initial step size of numerical solution is configured. (ms).

[0012] Preferably, the spacecraft dynamics library software currently configured in step S5 refers to the spacecraft dynamics library software whose orbital dynamics module, actuator module, and sensor module are optimized and configured according to step S3, and whose sensor module is configured with three-axis attitude equivalent measurement noise. rad, Gyroscope measurement noise rad / s, Furthermore, the actuator module is configured with a momentum wheel assembly to drive torque noise. Nm, momentum wheel assembly speed measurement noise rpm; The attitude dynamics module and numerical solution module are initially configured according to step S4; The integrated test environment used in step S5 refers to the closed-loop connection between the spacecraft dynamics library software and the GNC software under test according to the interaction interface between the spacecraft dynamics library software and the GNC software under test; In step S5, the parameters of the orbital dynamics module are configured according to the functions of the GNC software under test: orbital initial time (UTC). (s), initial position of the inertial frame (km), initial velocity of the inertial frame (km / s); Parameters of the attitude dynamics module: initial attitude angle (rad), initial attitude angular velocity (rad / s); Parameters of the numerical solution module: Spacecraft Dynamics Library Optimization Test Time Length (s); In step S5, the optimized configuration of the attitude dynamics module of the spacecraft dynamics library is determined through testing using the integrated test environment, as follows: (1) If the GNC software under test uses wheel control and attitude control, the optimized configuration of the attitude dynamics module adopts the initial configuration of the attitude dynamics module in step S4; perform one test using the integrated test environment: record the three-axis attitude angles in the attitude dynamics module. (rad), three-axis attitude angular velocity (rad / s); Record the position of the inertial frame in the orbital dynamics module. (km), velocity in inertial frame (km / s); where , , for and Multiples of the least common multiple of the function This indicates that the independent variable is rounded down to negative infinity; Record the total number of optimization tests for the attitude dynamics module. .

[0013] (2) If the GNC software under test uses jet attitude control, the steps to determine the optimal configuration of the attitude dynamics module through testing using the integrated test environment are as follows: Test 1: Configuration Record the three-axis attitude angles in the attitude dynamics module. (rad) and attitude angular velocity (rad / s); Record the position of the inertial frame in the orbital dynamics module. (km) and speed (km / s); where , , for and Multiples of the least common multiple; Second test: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , ;if , , Less than ,and Less than ;and Less than , Less than Then determine the attitude dynamics module configuration of the flexible spacecraft attitude dynamics and the minimum modal order. Otherwise, conduct a third test; Test 3: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , ;if , , Less than ,and Less than ;and Less than , Less than Then determine the attitude dynamics module configuration of the flexible spacecraft attitude dynamics and the minimum modal order. Otherwise, conduct a fourth test; Test 4: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , ;if , , Less than ,and Less than ;and Less than , Less than Then the attitude dynamics module is configured as flexible spacecraft attitude dynamics with the minimum modal order. Otherwise, determine that the attitude dynamics module is configured as flexible spacecraft attitude dynamics with the minimum modal order. ; The total number of times the attitude dynamics module optimization test was recorded was: .

[0014] Preferably, in step S6, the currently configured spacecraft dynamics library software refers to the spacecraft dynamics library software whose orbital dynamics module, actuator module, and sensor module are optimized and configured according to step S3, and whose sensor module is configured with three-axis attitude equivalent measurement noise. rad, Gyroscope measurement noise rad / s, Furthermore, the actuator module is configured with a momentum wheel assembly to drive torque noise. Nm, momentum wheel assembly speed measurement noise The numerical solution module is initially configured according to step S4; the attitude dynamics module is optimized according to step S5; in step S6, the integrated test environment used for testing refers to the closed-loop connection between the spacecraft dynamics library software and the GNC software under test according to the interaction interface between the spacecraft dynamics library software and the GNC software under test; in step S6, the parameters of the same orbital dynamics module are configured according to step S5: orbital initial time (UTC). (s), initial position of the inertial frame (km), initial velocity of the inertial frame (km / s); Parameters of the attitude dynamics module: initial attitude angle (rad), initial attitude angular velocity (rad / s); Parameters of the numerical solution module: Spacecraft Dynamics Library Optimization Test Time Length (s); In step S6, the optimized configuration of the numerical solution module of the spacecraft dynamics library is determined through testing using the integrated test environment, as follows: Test 1: Configuration (ms), record the three-axis attitude angles in the attitude dynamics module. (rad) and attitude angular velocity (rad / s); Record the position of the inertial frame in the orbital dynamics module. (km) and speed (km / s); where , and ;if , , Less than ,and Less than ;and Less than , Less than Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, conduct a second test; Second test: Configuration (ms), where Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , and ;if , , Less than ,and Less than ;and Less than , Less than Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, proceed to the next step. This test, among which ; No. Second test: Configuration (ms), where Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , and ;if , , Less than ,and Less than ;and Less than , Less than Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, proceed to the next step. Second test; No. Second test: Configuration (ms), record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , and ;if , , Less than ,and Less than ;and Less than , Less than Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, determine the maximum numerical solution step size of the numerical solution module. ; The total number of optimization tests recorded for the numerical solution module is .

[0015] The advantages of this invention compared to the prior art are: (1) The orbital dynamics module, attitude dynamics module, actuator module and sensor module of the spacecraft dynamics library constructed by the method of the present invention respectively realize the software simulation of the spacecraft orbit, attitude, actuator and sensor functions, and the numerical solution module realizes the numerical solution function of the spacecraft orbit dynamics equation and attitude dynamics equation, providing a comprehensive spacecraft dynamics library for realizing the GNC software integration test of high orbit spacecraft.

[0016] (2) The method of the present invention optimizes the configuration of perturbation terms in orbital dynamics, the configuration of attitude dynamics type (no attitude dynamics, rigid body attitude dynamics, flexible attitude dynamics), the configuration of modal order in flexible attitude dynamics, the configuration of external disturbance torque terms in attitude dynamics, and the configuration of numerical solution step size. This results in a spacecraft dynamics library with low computational complexity, effectively realizing a significant improvement in the integration and testing efficiency of high-orbit spacecraft GNC software, especially the long-term functionality of GNC software, without increasing the cost of hardware testing equipment.

[0017] (3) The interfaces between the functional modules of the spacecraft dynamics library constructed by the method of the present invention are fixed. Each functional module can be configured separately according to the function of the GNC software under test without affecting other modules. It has good scalability and is easy to implement in engineering.

[0018] In summary, the present invention provides an effective method for constructing a spacecraft dynamics library to improve the efficiency of GNC software integration testing for high-orbit spacecraft. The technology is easy to implement and has been verified in GNC software integration testing for high-orbit spacecraft. Attached Figure Description

[0019] Figure 1 This is a flowchart of the method for constructing a rapid integrated testing dynamics library for GNC software in high-orbit spacecraft according to the present invention; Figure 2 This is an integrated test environment consisting of the spacecraft dynamics library software of this invention and the GNC software under test; Figure 3 This invention utilizes two numerical solution step sizes in its embodiments. and The CPU time curve used for the integration test of the spacecraft dynamics library with the GNC software function under test (the spacecraft dynamics library optimization test time is limited to 1 day) is shown in (ms). Figure 4 This invention utilizes two numerical solution step sizes in its embodiments. and The satellite inertial frame position error curve (top figure) and velocity error curve (bottom figure) are obtained from the integration test of the spacecraft dynamics library with the tested GNC software function (the spacecraft dynamics library optimization test time is limited to 1 day). Figure 5 This embodiment of the invention utilizes configured numerical solution step size. The cumulative CPU time used by the spacecraft dynamics library for the integration test of the tested GNC software function (actual duration of the tested GNC software function is 30 days) in (ms). Figure 6 This embodiment of the invention utilizes configured numerical solution step size. The daily CPU time used by the spacecraft dynamics library for the integration test of the GNC software function under test (the actual duration of the GNC software function under test is 30 days) in (ms). Detailed Implementation

[0020] The following is in conjunction with the appendix Figure 1-6 The specific embodiments of the present invention will be described in further detail below.

[0021] like Figure 1 As shown, the method for constructing a rapid integrated test dynamics library for high-orbit spacecraft GNC software of the present invention specifically includes the following steps: Step S1: Design the functional modules of the spacecraft dynamics library; Step S2: Program the spacecraft dynamics library software. Step S3: Determine the optimized configuration of the orbital dynamics module, actuator module, and sensor module of the spacecraft dynamics library based on the functions of the GNC software under test; Step S4: Determine the initial configuration of the attitude dynamics module and numerical solution module of the spacecraft dynamics library based on the functions of the GNC software under test; Step S5: An integrated test environment is formed by using the currently configured spacecraft dynamics library software and the GNC software under test in a closed loop (e.g., ...). Figure 2 As shown in the figure, the optimal configuration of the attitude dynamics module of the spacecraft dynamics library was determined through testing; Step S6: The spacecraft dynamics library software with the current configuration and the GNC software under test are used to form an integrated test environment. The optimal configuration of the numerical solution module of the spacecraft dynamics library is determined through testing.

[0022] Among them, the functional modules of the spacecraft dynamics library in step S1 involve coordinate systems: J2000 geocentric inertial coordinate system (hereinafter referred to as inertial system), orbital coordinate system (hereinafter referred to as orbital system), spacecraft body coordinate system (hereinafter referred to as the body system), and Earth-fixed coordinate system (hereinafter referred to as Earth-fixed system). For the specific definitions of the four coordinate systems, please refer to the reference [1] (Li Hengnian, Geostationary Satellite Orbit and Co-location Control Technology, National Defense Industry Press, 2010). The transformation matrix between coordinate systems is involved: the transformation matrix from the inertial system to the body system. Transformation matrix from orbital system to this system Transformation matrix from inertial frame to Earth-fixed frame Transformation matrix from inertial frame to orbital frame The calculation methods for the four transformation matrices can be found in reference [1] (Li Hengnian, Geostationary Satellite Orbit and Co-location Control Technology, National Defense Industry Press, 2010). The functional module design steps of the spacecraft dynamics library specifically include: 2.1 Orbital dynamics module design, specifically including: (1) Orbital time Orbital Time (s), where The initial time of the orbit (UTC, in seconds). Let be the current simulation time (s) for orbital dynamics, and require... No earlier than the base time start point Configure according to requirements; requires configuring the base time starting point. No earlier than 12:00:00 UTC on January 1, 2000, in seconds; (2) Orbital dynamics equations Under the influence of Earth's gravity and external perturbations, the orbital dynamics equations of the spacecraft in the inertial frame are as follows: in, and These represent position and velocity in the inertial frame, respectively. The distance from the Earth's center to the spacecraft. km 3 / s 2 The gravitational constant of Earth; Earth's gravitational perturbation (km / s) 2 ), Solar gravitational perturbation (km / s) 2 ), Lunar gravitational perturbation (km / s) 2 ), Solar radiation pressure perturbation (km / s) 2 ), For orbital jet perturbation (km / s) 2 ); Step 2.1, the orbital dynamics module includes 5 perturbation terms, namely: Earth's gravitational perturbation Considering only the potential function of second-order Earth gravitational perturbation, i.e. ,in The geocentric longitude of the spacecraft. It is the geocentric latitude of the spacecraft. These are the position components of the spacecraft in the Earth-solid system, i.e. , km is the Earth's average equatorial radius, and the coefficient of the second-order zonal harmonic term. Second-order harmonic coefficient Geographic longitude of the second-order field harmonic term principal axis deg; Solar gravitational perturbation ,in km 3 / s 2 The gravitational constant of the Sun, This represents the vector (km) from the spacecraft to the sun with a magnitude of ; Represents the vector (km) from Earth to the Sun with a magnitude of ,in Given from the ephemeris; Lunar gravitational perturbation ,in km 3 / s 2 The gravitational constant of the moon, Represents the vector (km) from the spacecraft to the moon with a magnitude of ; Represents the vector (km) from Earth to the Moon with a magnitude of ,in Given from the ephemeris; Solar radiation pressure perturbation ,in For sun-receiving factors, Indicates the area of ​​the ground shadow. Indicates the area under illumination. The surface reflectance of the spacecraft is , and the solar radiation pressure intensity is . (N / m) 2 ), Solar radiation pressure surface mass ratio (m) 2 / kg), bound according to spacecraft configuration and material properties; Orbit-controlled jet perturbation The thrust vector of the orbital control engine in the body is Unit N, modulus is , The real-time mass (kg) of the entire satellite, and , , , This is the initial mass of the entire star. Indicates the rail control engine in the first The jet pulse width (s) for each control cycle is determined by the GNC software under test; and Indicates the first Specific impulse of individual attitude control thrusters This indicates the specific impulse of the rail control engine. Represents the acceleration due to gravity at the Earth's surface (m / s²). 2 ); (3) Orbital dynamics module configuration parameters, including: control period (Require (Unit: seconds); initial orbital time (UTC) (Configuration required no earlier than) (Unit: seconds); reference time start point (Configuration must be no earlier than 12:00:00 UTC on January 1, 2000), unit: seconds; reference time offset. Indicates the starting point of the reference time. The number of seconds relative to 12:00:00 UTC on January 1, 2000 (configuration required). ), unit s; sun exposure factor Take in the shadow area Take in the lighted area Spacecraft surface reflectivity Typically, 1.5 is used; solar radiation pressure surface mass ratio Unit m 2 / kg; initial mass of the spacecraft Unit: kg; Thrust vector of the orbital control engine in this system. Unit: N; Specific impulse of the rail control engine Unit: s; Initial position of the inertial frame. Unit: km; Initial velocity of the inertial frame. Unit: km / s; Track dynamics module storage period ( (positive integer), unit s; allowable error limit of inertial frame position. Unit: km; Permissible error limit of inertial frame velocity The unit is km / s; (4) Variables recorded by the orbital dynamics module, including: the number of orbital times. The unit is s, where The number of positions in the inertial frame. Unit: km; Number of inertial frame velocities The unit is km / s; 2.2 Attitude dynamics module design, specifically including: (1) Posture kinematic equations ( , and Values (But they are all different) The kinematic equations of the spacecraft attitude under the transition sequence are: in, These represent the roll angle, pitch angle, and yaw angle of the spacecraft relative to its orbital system, respectively. This represents the component of the spacecraft's instantaneous rotational speed relative to the orbital system within this system. Represents Euler angular rate and Relationship matrix, 6 different transformations The expression for is found in Chapter 2 of reference [2] (Tu Shancheng, ed. Satellite Attitude Dynamics and Control. Aerospace Press. 2001); , The rotational velocity of the body relative to the inertial frame is expressed in this system. The rotational speed of the orbital frame relative to the inertial frame is expressed in the orbital frame, where the Earth's gravitational constant is... km 3 / s 2 , The semi-major axis of the spacecraft's orbit (km). This represents the transformation matrix from the orbital system to the home system, while Indicate Euler angles at The attitude matrix corresponding to the transposed order. For the specific expression, please refer to Chapter 2 of reference [2] (Tu Shancheng, ed. Satellite Attitude Dynamics and Control. Aerospace Press. 2001).

[0023] (2) Attitude dynamics equations Flexible spacecraft dynamics (with symmetrical solar panels, considering only the effect of panel vibration on the rotation of the central rigid body, ignoring the effect of panel vibration on the translation and rotation of the central rigid body, and ignoring the effect of panel rotation on the rotation of the central rigid body) has two configurable types, as follows: (i) Momentum wheel control (referred to as wheel-controlled attitude control) attitude dynamics of flexible spacecraft in, For the entire star's inertia matrix (relative to the center of mass); The installation matrix of the momentum wheel assembly. ; , The moment of inertia matrix of the momentum wheel assembly, in kg. m 2 ,in The number of momentum wheels, where For the first The moment of inertia of a momentum wheel; Let be the speed vector of the momentum wheel assembly, in rad / s, where For the first The rotational speed of the momentum wheel; and This represents the flexible coupling coefficient matrix of the vibrations of the left and right solar panels on the rotation of the spacecraft (central rigid body). ,in The modal order; and These are the modal coordinate vectors for the left and right solar panels; and This is the diagonal matrix of the modal frequencies of the left and right solar panels; and This is a diagonal matrix of the modal damping coefficients for the left and right solar panels, with the diagonal elements typically set to 0.005. Indicates the wheel control attitude control torque; The driving torque vector of the momentum wheel assembly (calculated by the GNC software under test), in Nm; The driving torque noise of the momentum wheel assembly, where This indicates that the mean is 0 and the variance is... Noise level, in Nm; This is the diagonal matrix in the direction of the momentum wheel assembly's rotational speed. ; This is the set value of the dynamic friction torque of the momentum wheel assembly, in Nm; The damping friction coefficient of the momentum wheel is given in Nm / rpm; the nominal speed vector of the momentum wheel assembly is given. (Configurable), unit rpm; the operating speed vector range of the momentum wheel assembly is... (Configurable), where , The unit is rpm; Represents the external disturbance torque, where The solar pressure torque is expressed in triaxial form within this system, in Nm. ,and , and , This is the optical pressure torque coefficient, which is related to the geometry and mass distribution characteristics of the spacecraft. The value is the representation of the track control interference torque in this system, and is a bound value in Nm; (ii) Jet control (or jet attitude control) attitude dynamics of flexible spacecraft Among them, the The lever arm vector of the attitude control thruster The thrust vectors of the three axes in this system , Indicates the first The jet pulse sequence of an attitude control thruster is defined as follows: The assignment logic of the attitude control thruster command in the tested GNC software is determined. Indicates the first The attitude control thruster in the first Jet pulse width per control cycle, in seconds. ; (3) Attitude dynamics module configuration parameters, including: sequence transition It includes 6 types of sequence transformations: , , , , , Inertia matrix of the entire star (Relative center of mass), unit kg.m 2 , Flexible coupling coefficient matrix of left and right solar panel vibrations to the rotation of the central rigid body and , Modal order Modal frequencies of the left and right solar panels and Diagonal matrix of modal damping coefficients for left and right solar panels and The diagonal elements are generally taken as 0.005; the number of momentum wheels Moment of inertia of the momentum wheel , Unit: kg m 2 ; Moment of inertia matrix of momentum wheel assembly , Unit: kg m 2 Installation matrix of momentum wheel assembly dimensionality Momentum wheel assembly driving torque noise Unit: Nm; Momentum wheel assembly speed measurement noise Unit: rpm; dynamic friction torque of momentum wheel assembly (set value) The unit is Nm; the damping friction coefficient of the momentum wheel. Unit: Nm / rpm; Number of attitude control thrusters Attitude control thruster nozzle installation position , , unit m; attitude control thruster nozzle vector angle (angle with the positive directions of the three axes of the body). , Unit: deg; Nominal force of attitude control thruster , Unit N; the first The specific impulse of the individual attitude control thruster , Unit: seconds; Initial attitude angle , unit: rad; initial attitude angular velocity Unit: rad / s; Attitude dynamics module data storage cycle ( (positive integer), unit s; attitude angle tolerance limit Unit: rad; allowable error limit for attitude angular velocity. The unit is rad / s; (4) Variables recorded by the attitude dynamics module, including: the stored values ​​of the three-axis attitude angles. ,in , in rad, where Storage of three-axis attitude angular velocities The unit is rad / s, where Total number of optimization tests for attitude dynamics module ; 2.3 Actuator module design, specifically including: (1) Momentum wheel design: based on Calculate The rotational speed of the momentum wheel assembly composed of momentum wheels (rad / s), then we get Measuring the rotational speed of the momentum wheel (rpm), output to the GNC software under test; where For momentum wheel assembly speed measurement noise, The mean is 0 and the variance is Noise (rpm) ; (2) Attitude control thruster design: according to Calculate the attitude control torque (Nm) acting on the three axes of the spacecraft body during the current control cycle, and then drive the spacecraft attitude evolution; (3) Track control engine design: according to Calculate the orbital control force (N) applied to the three axes of the spacecraft's inertial frame during the current control cycle, and then drive the spacecraft's orbital evolution; 2.4 Sensor module design, specifically including: (1) Star sensor design: For star sensors The attitude dynamics module outputs three-axis attitude angles. Unit: rad, with added three-axis attitude equivalent measurement noise. The unit is rad. ,according to The attitude matrix of this system relative to the orbital system is obtained by transposing the order. Then according to Calculation of star sensitivity Measured spacecraft attitude matrix Convert to the corresponding quaternion Then we get Xingmin Measured spacecraft attitude quaternions ;in , and The mean is 0 and the variance is noise, Xingmin Installation matrix (configurable), number of star sensors (Configurable) ; (2) Gyroscope design: according to Calculate the first The angle increment of the gyroscope only , in rad; where The three-axis attitude angular velocity output by the attitude dynamics module, in rad / s; For the first Add a constant drift value to the gyroscope, in rad / s; With a mean of 0 and a variance of The random drift, in rad / s; Indicates the first Mounting vector for only the gyroscope input axis. To control the period, the number of gyroscopes (Configurable), and the configuration must ensure that at least 3 gyroscopes are not coplanar. ; (3) Sensor module configuration parameters, including: number of star sensors Starsense Installation Matrix , Three-axis attitude equivalent measurement noise , Number of spinning tops ; Gyroscope mounting vector ,in ; Gyroscope constant drift ,in ; Gyroscope measurement noise ,in ; 2.5 Numerical solution module design, specifically including: (1) Numerical solution algorithm, using the fourth-order Runge-Kutta algorithm to realize the numerical solution of the spacecraft attitude dynamics equation and orbital dynamics equation; (2) Numerical solution step size The unit is ms, which is the configuration parameter to be optimized; (3) Configuration parameters of the numerical solution module, including: numerical solution step size Unit: milliseconds; Dynamics library optimization test time. That is, optimizing the time interval of the dynamics library through testing. ,in Based on the configuration of the tested GNC software, the unit is seconds; the actual duration of the tested GNC software function. That is, the actual time interval of the tested GNC software function is ,in According to the functional configuration of the tested GNC software, the unit is s; (4) Variables recorded by the numerical solution module, including: the total number of optimization tests of the numerical solution module. ; 2.6 Interface design between functional modules of the spacecraft dynamics library, including: (1) Input of the orbital dynamics module, including: the current state of the orbital dynamics equation (output of the numerical solution module), and the orbital control force (output of the actuator module); (2) Output of the orbital dynamics module, including: the state of the orbital dynamics equation of the previous long time step (for the numerical solution module); (3) Input of the attitude dynamics module, including: the current state of the attitude dynamics equation (output of the numerical solution module), and the attitude control torque (output of the actuator module); (4) Output of the attitude dynamics module, including: the state of the attitude dynamics equation of the previous long time step (for the numerical solution module), and the spacecraft attitude angle and attitude angular velocity. (5) Outputs of the actuator module, including: attitude control torque (to attitude dynamics module), orbit control force (to orbit dynamics module); (6) Inputs of the sensor module, including: spacecraft attitude angle and attitude angular velocity (outputs of attitude dynamics module); (7) Inputs of the numerical solution module, including: attitude dynamics equation state of the previous long time step (output of attitude dynamics module), orbit dynamics equation state of the previous long time step (output of orbit dynamics module); (8) Outputs of the numerical solution module, including: attitude dynamics equation state of the current time step (to attitude dynamics module), orbit dynamics equation state of the current time step (to orbit dynamics module).

[0024] Among them, the spacecraft dynamics library software obtained by programming in step S2 refers to the software obtained by using a programming language consistent with the programming language of the GNC software under test, according to the five functional modules in steps 2.1 to 2.5 and the interface between the five functional modules of the dynamics library in step 2.6; the interaction interface between the spacecraft dynamics library software and the GNC software under test includes: (1) the orbital time output by the orbital dynamics module. Spacecraft position and speed (GNC software under test); (2) Quaternion of spacecraft attitude measured by star sensor output by sensor module Angular increment measured by gyroscope (GNC software under test); (3) The momentum wheel output by the actuator module measures the rotational speed. (GNC software under test); (4) Wheel control attitude control torque input to the actuator module Attitude control jet pulse width Orbit control jet pulse width (Output from the tested GNC software).

[0025] The optimized configuration of the orbital dynamics module, actuator module, and sensor module of the spacecraft dynamics library in step S3 is as follows: (1) Optimized configuration of orbital dynamics module: The orbital dynamics module is configured with Earth's central gravitational acceleration. If the tested GNC software function involves orbit control jet thrust, then the orbit dynamics module should be configured with the orbit control jet thrust item. Otherwise, do not configure. If the tested GNC software function involves longitude... Peaceful longitude change rate Then the orbital dynamics module is configured with Earth's gravitational perturbation terms. Otherwise, do not configure. If the GNC software function being tested involves tilt angle... Right ascension of ascending node Then the orbital dynamics module is configured with solar gravitational perturbation terms. and lunar gravitational perturbation terms Otherwise, do not configure. and If the tested GNC software function involves eccentricity... Perigeal argument Then the orbital dynamics module is configured with solar radiation pressure perturbation terms. Otherwise, do not configure. ; (2) Optimized configuration of actuator module: Determine the configuration parameters of the actuator module based on the functions of the GNC software under test: number of momentum wheels Moment of inertia of the momentum wheel , Unit: kg m 2 ; Moment of inertia matrix of momentum wheel assembly , Unit: kg m 2 Installation matrix of momentum wheel assembly dimensionality Momentum wheel assembly driving torque noise Unit: Nm; Momentum wheel assembly speed measurement noise Unit: rpm; dynamic friction torque of momentum wheel assembly (set value) The unit is Nm; the damping friction coefficient of the momentum wheel. Unit: Nm / rpm; Number of attitude control thrusters Attitude control thruster nozzle installation position , , unit m; attitude control thruster nozzle vector angle (angle with the positive directions of the three axes of the body). , Unit: deg; Nominal force of attitude control thruster , Unit N; the first The specific impulse of the individual attitude control thruster , Unit: s; Thrust vector of the orbital control engine in this system. Unit: N; Specific impulse of the rail control engine Unit: s; (3) Sensor module optimization configuration: Determine the sensor module configuration parameters based on the functions of the GNC software under test: number of star sensors Starsense Installation Matrix , Three-axis attitude equivalent measurement noise , Number of spinning tops ; Gyroscope mounting vector ,in ; Gyroscope constant drift ,in ; Gyroscope measurement noise ,in .

[0026] The initial configuration steps for the attitude dynamics module and numerical solution module of the spacecraft dynamics library in step S4 are as follows: (1) Initial configuration of attitude dynamics module: If the GNC software under test involves spacecraft position and velocity, but not spacecraft attitude angle and attitude angular velocity, and wheel control is used for attitude control, then the attitude dynamics module is configured directly. rad, rad / s, meaning the attitude dynamics module no longer performs numerical calculations; If the GNC software under test involves spacecraft attitude angles and attitude angular velocities, and uses wheel-based attitude control without orbit control, then rigid body spacecraft attitude dynamics should be configured (i.e., configuration of...). and ), wheel-controlled attitude control torque and solar radiation pressure torque; that is, the spacecraft attitude dynamics equation is If the GNC software function being tested involves spacecraft attitude angles and attitude angular velocities, and uses wheel-based attitude control and jet orbit control, then rigid body spacecraft attitude dynamics should be configured (i.e., configured). and The spacecraft attitude dynamics equations are: wheel control attitude control torque, solar radiation pressure torque, and orbital control disturbance torque; that is, the spacecraft attitude dynamics equations are: If the tested GNC software function uses jet attitude control and has no orbit control, then configure the flexible spacecraft attitude dynamics and jet attitude control torque, including configuring the modal order. External disturbance torque That is, the spacecraft attitude dynamics equation is If the tested GNC software function uses jet attitude control and jet orbit control, then configure the flexible spacecraft attitude dynamics, jet attitude control torque, and orbit control disturbance torque, including configuring the modal order. External disturbance torque That is, the spacecraft attitude dynamics equation is (2) Initial configuration of numerical solution module: The numerical solution algorithm adopts the fourth-order Runge-Kutta algorithm, and the initial step size of numerical solution is configured. (ms).

[0027] In step S5, the spacecraft dynamics library software currently configured refers to the spacecraft dynamics library software obtained according to step S2, wherein the orbital dynamics module, actuator module, and sensor module are optimized and configured according to step S3, and the sensor module is configured with three-axis attitude equivalent measurement noise. rad, Gyroscope measurement noise rad / s, Furthermore, the actuator module is configured with a momentum wheel assembly to drive torque noise. Nm, momentum wheel assembly speed measurement noise rpm; The attitude dynamics module and numerical solution module are initially configured according to step S4; The integrated test environment used in step S5 refers to the closed-loop connection between the spacecraft dynamics library software and the GNC software under test according to the interaction interface between the spacecraft dynamics library software and the GNC software under test; In step S5, the parameters of the orbital dynamics module are configured according to the functions of the GNC software under test: orbital initial time (UTC). (s), initial position of the inertial frame (km), initial velocity of the inertial frame (km / s); Parameters of the attitude dynamics module: initial attitude angle (rad), initial attitude angular velocity (rad / s); Parameters of the numerical solution module: Spacecraft Dynamics Library Optimization Test Time Length (s); In step S5, the optimized configuration of the attitude dynamics module of the spacecraft dynamics library is determined through testing using the integrated test environment, as follows: (1) If the GNC software under test uses wheel control and attitude control, the optimized configuration of the attitude dynamics module is the initial configuration of the attitude dynamics module in step S4; perform one test using the integrated test environment: record the three-axis attitude angles in the attitude dynamics module. (rad), three-axis attitude angular velocity (rad / s); Record the position of the inertial frame in the orbital dynamics module. (km), velocity in inertial frame (km / s); where , , for and Multiples of the least common multiple of the function This indicates that the independent variable is rounded down to negative infinity; Record the total number of optimization tests for the attitude dynamics module. .

[0028] (2) If the GNC software under test uses jet attitude control, the steps to determine the optimal configuration of the attitude dynamics module through testing using the integrated test environment are as follows: Test 1: Configuration Record the three-axis attitude angles in the attitude dynamics module. (rad) and attitude angular velocity (rad / s); Record the position of the inertial frame in the orbital dynamics module. (km) and speed (km / s); where , , for and Multiples of the least common multiple; Second test: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , ;if , , Less than ,and Less than ;and Less than , Less than Then determine the attitude dynamics module configuration of the flexible spacecraft attitude dynamics and the minimum modal order. Otherwise, conduct a third test; Test 3: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , ;if , , Less than ,and Less than ;and Less than , Less than Then determine the attitude dynamics module configuration of the flexible spacecraft attitude dynamics and the minimum modal order. Otherwise, conduct a fourth test; Test 4: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , ;if , , Less than ,and Less than ;and Less than , Less than Then the attitude dynamics module is configured as flexible spacecraft attitude dynamics with the minimum modal order. Otherwise, determine that the attitude dynamics module is configured as flexible spacecraft attitude dynamics with the minimum modal order. ; The total number of times the attitude dynamics module optimization test was recorded was: .

[0029] In step S6, the currently configured spacecraft dynamics library software refers to the spacecraft dynamics library software obtained according to step S2, wherein the orbital dynamics module, actuator module, and sensor module are optimized and configured according to step S3, and the sensor module is configured with three-axis attitude equivalent measurement noise. rad, Gyroscope measurement noise rad / s, Furthermore, the actuator module is configured with a momentum wheel assembly to drive torque noise. Nm, momentum wheel assembly speed measurement noise The numerical solution module is initially configured according to step S4; the attitude dynamics module is optimized according to step S5; in step S6, the integrated test environment used for testing refers to the closed-loop connection between the spacecraft dynamics library software and the GNC software under test according to the interaction interface between the spacecraft dynamics library software and the GNC software under test; in step S6, the parameters of the orbital dynamics module are configured according to the same parameters as in step S5: orbital initial time (UTC). (s), initial position of the inertial frame (km), initial velocity (km / s); Parameters of the attitude dynamics module: initial attitude angle (rad), initial attitude angular velocity (rad / s); Parameters of the numerical solution module: Spacecraft Dynamics Library Optimization Test Time Length (s); In step S6, the optimized configuration of the numerical solution module of the spacecraft dynamics library is determined through testing using the integrated test environment, as follows: Test 1: Configuration (ms), record the three-axis attitude angles in the attitude dynamics module. (rad) and attitude angular velocity (rad / s); Record the position of the inertial frame in the orbital dynamics module. (km) and speed (km / s); where , and ;if , , Less than ,and Less than ;and Less than , Less than Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, conduct a second test; Second test: Configuration (ms), where Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , and ;if , , Less than ,and Less than ;and Less than , Less than Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, proceed to the next step. This test, among which ; No. Second test: Configuration (ms), where Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , and ;if , , Less than ,and Less than ;and Less than , Less than Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, proceed to the next step. Second test; No. Second test: Configuration (ms), record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , and ;if , , Less than ,and Less than ;and Less than , Less than Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, determine the maximum numerical solution step size of the numerical solution module. ; The total number of optimization tests recorded for the numerical solution module is .

[0030] Example Select the onboard GNC software function to be measured: the longitude control function of a GEO satellite at geographic longitude 0 deg for 30 days (during which the satellite is in a free drift phase and has not triggered jet position holding control), and observe the satellite's geographic longitude. The changes.

[0031] (1) Design and optimization configuration of orbit dynamics module (i) Track timing design Orbital Time (s), where The initial orbital time (UTC) is to be configured, in seconds. This represents the current simulation time (in seconds) for orbital dynamics.

[0032] (ii) Design and optimization of orbital dynamics equations Since the tested GNC software only involves changes in the satellite's mean longitude, the orbital dynamics module is only configured with Earth's gravitational perturbation terms. That is, the orbital dynamics equations in the inertial frame are: in, and These represent the satellite's inertial frame position and velocity, respectively. The distance between the satellite and the Earth's center. km 3 / s 2 Earth's gravitational constant; Earth's gravitational perturbation (km / s) 2 ), including the Earth's gravitational perturbation potential function ,and It is the geocentric longitude of the satellite. It is the satellite's geocentric latitude; among which The position of the satellite in the Earth-fixed system, i.e. ; km is the Earth's average equatorial radius; second-order zonal harmonic coefficient. Second-order harmonic coefficient Geographic longitude of the second-order field harmonic term principal axis (deg); about For the calculation, please refer to reference [1] (Li Hengnian, Geostationary Satellite Orbit and Co-position Control Technology, National Defense Industry Press, 2010).

[0033] (iii) Configuration parameters of the orbital dynamics module Configure the control cycle according to the functions of the GNC software under test. (s), initial orbital time The reference time is 12:00:00 UTC on January 1, 2000. The initial position of the inertial frame is given by UTC time 12:00:00 on January 1, 2000 (s). (km), initial velocity (km / s), the orbital dynamics module's storage period is (s), i.e., configuration Inertial frame position tolerance limit (km), permissible error limit of inertial frame velocity (km / s).

[0034] (iv) Variables recorded by the orbital dynamics module Storage of orbital time (s), the number of positions in the inertial frame (km); the number of inertial frame velocities (km / s), where ; (2) Attitude dynamics module design and optimized configuration (i) Design and optimization of kinematic equations and attitude dynamics equations Since the tested GNC software functions involve satellite position and velocity, but not satellite attitude angle and attitude angular velocity, and uses wheel-based attitude control, the attitude dynamics module is directly configured. rad, rad / s, where the semi-major axis of the orbit (km), meaning the attitude dynamics module no longer performs numerical calculations; (ii) Attitude Dynamics Module Configuration Parameters Configure sequence 312, initial attitude angle (rad), initial attitude angular velocity (rad / s), the attitude dynamics module's data storage period is (s), i.e., configuration Attitude angle tolerance limit (rad), attitude angular velocity tolerance limit (rad / s); (iii) Variables recorded by the attitude dynamics module Storage of three-axis attitude angles (rad), the stored value of the three-axis attitude angular velocity (rad / s), where , ; (3) Actuator module design and optimized configuration (i) Attitude control thruster optimization configuration: Since the attitude dynamics module no longer performs numerical calculations, the momentum wheel optimization configuration is not involved; (ii) Attitude control thruster optimization configuration is not involved; (iii) Orbit control engine optimization configuration is not involved. (4) Sensor module design and optimized configuration (i) Optimized configuration of star sensor For Xingmin The attitude dynamics module outputs three-axis attitude angles. Unit: rad, with added three-axis attitude equivalent measurement noise. The unit is rad. ,according to The attitude matrix of this system relative to the orbital system is obtained by transposing the order. Then according to Calculation of star sensitivity Measured spacecraft attitude matrix Convert to the corresponding quaternion Then we get Xingmin Measured spacecraft attitude quaternions ;in , and The mean is 0 and the variance is noise, Xingmin Installation matrix (configurable), number of star sensors (Configurable) ; Configure the number of star sensors The mounting matrix of the star sensor is an identity matrix. Three-axis attitude equivalent measurement noise (rad); the three-axis attitude angles output by the attitude dynamics module. (rad), add three-axis attitude equivalent measurement noise get The attitude matrix of this system relative to the orbital system is obtained according to the 312 transformation sequence. According to the satellite position in the inertial frame output by the orbital dynamics module ,speed Calculated ,in , , Press again The attitude matrix of the star-sensor measurement frame relative to the inertial frame was calculated. Then convert it into the corresponding quaternion. That is, the quaternion of satellite attitude measured by star-sensor is ; (ii) Gyroscope optimization configuration Configure the number of gyroscopes The mounting vectors of each gyroscope are respectively , , gyroscope constant drift Gyroscope measurement noise ,in ;according to Calculation yields the first The angle increment of the gyroscope only (rad), of which This is the three-axis attitude angular velocity (rad / s) output by the attitude dynamics module. For control period (s); (5) Design and initial configuration of the numerical solution module: (i) The numerical solution algorithm adopts the fourth-order Runge-Kutta algorithm; (ii) Initial step size of the numerical solution. (ms); (iii) Spacecraft dynamics library optimization test duration (s), that is, the time interval for spacecraft dynamics library optimization testing is limited to... (s); (6) Design of functional module interfaces for spacecraft dynamics library (i) Input to the orbital dynamics module: the current state of the orbital dynamics equations and (Output from the numerical solution module), where (s), (ii) Output of the orbital dynamics module: the state of the orbital dynamics equations at the previous long time step. and (iii) Input to the attitude dynamics module, including: the current state of the attitude dynamics equations. and (Numerical solution module output), wheel control attitude control torque (iv) Output of the attitude dynamics module: satellite attitude angles and attitude angular velocity (v) Output of the actuator module: wheel control attitude control torque (to the sensor module); (v) Output of the actuator module: wheel control attitude control torque (For attitude dynamics module), orbit control force And orbital control jet pulse width , (For the orbital dynamics module); (vi) Input to the sensor module: satellite attitude angles and attitude angular velocity (Attitude Dynamics Module Output); (vii) Input to the Numerical Solution Module: The state of the orbital dynamics equations at the previous long time step. and (viii) Output of the orbital dynamics module; (viii) Output of the numerical solution module: Current state of the orbital dynamics equations. and (To the orbital dynamics module).

[0035] (7) Programmatically implement the spacecraft dynamics library software The orbital dynamics module optimized in step (1) of the embodiment, the attitude dynamics module optimized in step (2) of the embodiment, the actuator module optimized in step (3) of the embodiment, the sensor module optimized in step (4) of the embodiment, and the numerical solution module initially configured in step (5) of the embodiment are implemented in C language to obtain the spacecraft dynamics library software according to the spacecraft dynamics library functional module interface designed in step (6) of the embodiment.

[0036] (8) Determine the optimal configuration of the attitude dynamics module through testing. (8.1) Assemble the integrated test environment According to the interaction interface between the spacecraft dynamics library software and the GNC software under test: (i) the orbital time output by the orbital dynamics module Satellite inertial frame position and inertial frame velocity (i) The satellite attitude quaternion measured by the star sensor output by the sensor module; (ii) The satellite attitude quaternion measured by the star sensor. Angular increment measured by gyroscope (to the GNC software under test); (iii) the speed of the momentum wheel output by the actuator module. (iv) The wheel control torque input to the actuator module. Attitude control jet pulse width And orbital control jet pulse width ,in (Output from the GNC software under test) The spacecraft dynamics library software and the GNC software under test are connected in a closed loop to form an integrated test system.

[0037] (8.2) Configure initial values ​​for the spacecraft dynamics library and optimize test time Configure the parameters of the orbital dynamics module according to step (1) of the embodiment: orbital initial time (UTC) The initial position of the inertial frame is given by UTC time 12:00:00 on January 1, 2000 (s). (km), initial velocity of the inertial frame (km / s); Configure the parameters of the attitude dynamics module according to step (2) of the embodiment: initial three-axis attitude angles (rad), initial three-axis attitude angular velocity (rad / s); Configure the parameters of the numerical solution module according to step (5) of the embodiment: Spacecraft Dynamics Library Optimization Test Time Length (s); (8.3) Record test data Record the three-axis attitude angles in the attitude dynamics module. (rad), three-axis attitude angular velocity (rad / s); Record the position of the inertial frame in the orbital dynamics module. (km), speed (km / s); where , (Right now Record the total number of optimization tests for the attitude dynamics module. .

[0038] (9) Determine the optimal configuration of the numerical solution module through testing. (9.1) The composition of the integrated testing environment is the same as the steps in the embodiment (8.1); (9.2) The initial values ​​of the spacecraft dynamics library and the configuration of the optimized test time are the same as the steps in the embodiment (8.2); (9.3) Configuration (ms), record test data Record the three-axis attitude angles in the attitude dynamics module. (rad), three-axis attitude angular velocity (rad / s); Record the position of the inertial frame in the orbital dynamics module. (km), velocity in inertial frame (km / s); where , ; (9.4) Determine the maximum numerical solution step size using test data. Using the test data from steps (9.3) and (8.3) of the embodiment, the three-axis attitude angle errors were obtained as follows: , , The three-axis attitude angular velocity error is The position error in the inertial frame is The velocity error in the inertial frame is ,in , Then determine the maximum numerical solution step size of the numerical solution module. The total number of optimization tests recorded for the numerical solution module is [number]. .

[0039] The functionality of the GNC software under test was tested using an optimized spacecraft dynamics library software: The optimized configuration of the spacecraft dynamics library software refers to the following: the orbital dynamics module, actuator module, and sensor module of the spacecraft dynamics library software are optimized using steps (1), (3), and (4) respectively, and the storage period of the orbital dynamics module is [missing information]. ( It is a positive integer. The attitude dynamics module adopts the optimized configuration of step (2) and the storage period of the attitude dynamics module is [missing information]. ( (where the integer is positive); the numerical solution module adopts the optimized configuration of step (9); The integrated test environment used in the test refers to: closing the loop between the spacecraft dynamics library software and the GNC software under test according to the interaction interface between the spacecraft dynamics library software and the GNC software under test; Configure the parameters of the orbit dynamics module according to the functions of the GNC software under test: orbit initial moment The initial position of the satellite's inertial frame at 12:00:00 UTC on January 1, 2000 (s). (km), the initial velocity of the satellite's inertial frame. (km / s), the storage period of the orbital dynamics module is (s), i.e., configuration Parameters of the attitude dynamics module: initial three-axis attitude angles (rad), initial attitude angular velocity (rad / s), the storage period of the attitude dynamics module is (s), i.e., configuration Parameters of the numerical solution module: Actual duration of the tested GNC software function. (s), that is, the actual time interval of the tested GNC software function. (s); Record the test data corresponding to the functions of the GNC software under test: record the three-axis attitude angles in the attitude dynamics module. (rad), three-axis attitude angular velocity (rad / s); Record the position of the inertial frame in the orbital dynamics module. (km), speed (km / s) and satellite geographic longitude (deg); where , (Right now ).

[0040] Based on the test results, the following conclusions are drawn: 1. For example Figure 3 As shown, two numerical methods are used to solve for the step size. and (ms) CPU time used for the integration test of the GNC software function under test (test duration limited to 1 day) by the spacecraft dynamics library configuration, from Figure 3 It can be seen that increasing the numerical solution step size can significantly reduce the CPU time consumed in integration testing and effectively improve the efficiency of integration testing.

[0041] 2. For example Figure 4 As shown, two numerical methods are used to solve for the step size. and The spacecraft dynamics library configured (ms) integrates the tested GNC software functions (test duration limited to 1 day) to test the satellite inertial frame position and velocity errors. Figure 4 As can be seen, the method of the present invention significantly reduces the CPU time consumption for integrated testing without affecting the test accuracy.

[0042] 3. For example Figure 5 As shown, the step size is solved numerically. The cumulative CPU time used by the spacecraft dynamics library configured in (ms) for the integration test of the GNC software function under test (the actual duration of the GNC software function under test is 30 days), from Figure 5 It can be seen that the CPU time consumed in the integration test increases linearly with the duration of the tested GNC software function.

[0043] 4. For example Figure 6 As shown, the step size is solved numerically. The daily CPU time used by the spacecraft dynamics library configured (ms) for the integration test of the GNC software function under test (actual duration of the GNC software function under test is 30 days) is calculated from... Figure 6 It can be seen that the CPU time consumed during the daily integration test of the tested GNC software functions is consistently around 300 seconds.

[0044] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for constructing a rapid integrated test dynamics library for GNC software of high-orbit spacecraft, characterized in that: include: Design functional modules for the spacecraft dynamics library; The spacecraft dynamics library software was obtained through programming. Based on the functions of the tested GNC software, determine the optimal configuration of the orbital dynamics module, actuator module, and sensor module of the spacecraft dynamics library; Based on the functions of the GNC software under test, determine the initial configuration of the attitude dynamics module and numerical solution module of the spacecraft dynamics library; An integrated test environment was formed by using the currently configured spacecraft dynamics library software and the GNC software under test in a closed loop. The optimal configuration of the attitude dynamics module of the spacecraft dynamics library was determined through testing. An integrated test environment was formed by using the currently configured spacecraft dynamics library software and the GNC software under test in a closed loop. The optimal configuration of the numerical solution module of the spacecraft dynamics library was determined through testing. The functional modules of the spacecraft dynamics library The coordinate systems involved are: J2000 geocentric inertial coordinate system (hereinafter referred to as the inertial system); orbital coordinate system (hereinafter referred to as the orbital system); spacecraft body coordinate system (hereinafter referred to as the body system); and Earth-fixed coordinate system (hereinafter referred to as the Earth-fixed system). The transformation matrix involves four types of coordinate systems: the transformation matrix from the inertial frame to the local frame. Transformation matrix from orbital system to this system Transformation matrix from inertial frame to Earth-fixed frame The transformation matrix from inertial frame to orbital frame ; The functional module design steps of the spacecraft dynamics library specifically include: 3.1 Orbital dynamics module design, specifically including: (3.11) Orbital Time Orbital Time ,in The initial time of the orbit is UTC. The current simulation time for orbital dynamics is given, and it is required that... No earlier than the reference time start point Configure according to requirements; requires configuring the base time starting point. No earlier than 12:00:00 UTC on January 1, 2000; (3.12) Orbital dynamics equations Under the influence of Earth's gravity and external perturbations, the orbital dynamics equations of the spacecraft in the inertial frame are as follows: in, and These represent position and velocity in the inertial frame, respectively. The distance from the Earth's center to the spacecraft. km 3 / s 2 The gravitational constant of Earth; Perturbation by Earth's gravity perturbed by the sun's gravity, For lunar gravitational perturbation, For solar radiation pressure perturbation, For orbital control jet perturbation; The orbital dynamics module includes five perturbation terms, namely: Earth's gravitational perturbation Considering only the potential function of second-order Earth gravitational perturbation, i.e. ,in It is the geocentric longitude of the spacecraft. It is the geocentric latitude of the spacecraft. These are the position components of the spacecraft in the Earth-solid system, i.e. , km is the Earth's average equatorial radius, and the coefficient of the second-order zonal harmonic term. Second-order harmonic coefficient Geographic longitude of the second-order field harmonic term principal axis deg; Solar gravitational perturbation ,in km 3 / s 2 The gravitational constant of the Sun, Represents the vector from the spacecraft to the sun with a magnitude of ; Represents the vector from Earth to the Sun with a magnitude of ,in Given from the ephemeris; Lunar gravitational perturbation ,in km 3 / s 2 The gravitational constant of the moon, Represents the vector from the spacecraft to the moon with a magnitude of ; Represents the vector from Earth to the Moon with a magnitude of ,in Given from the ephemeris; Solar radiation pressure perturbation ,in For sun-receiving factors, Indicates the area of ​​the ground shadow. Indicates the area under illumination. The surface reflectance of the spacecraft is , and the solar radiation pressure intensity is . , The solar radiation pressure-to-mass ratio is determined and bound according to the spacecraft configuration and material properties. Orbit-controlled jet perturbation The thrust vector of the orbital control engine in the body is The modulus is , For the real-time quality of the entire satellite, and , , , This is the initial mass of the entire star. Indicates the rail control engine in the first The jet pulse width for each control cycle is determined by the GNC software under test; and Indicates the first Specific impulse of individual attitude control thrusters This indicates the specific impulse of the rail control engine. This represents the acceleration due to gravity at the Earth's surface. (3.13) Orbital dynamics module configuration parameters, including: control period ,Require Initial moment of the orbit Configuration required no earlier than ; Baseline time start point The configuration must be no earlier than 12:00:00 UTC on January 1, 2000; reference time offset. Indicates the starting point of the reference time. The required configuration is as follows: The number of seconds between the time interval and the time of January 1, 2000, 12:00:00 UTC. Sun exposure factor Take in the shadow area Take in the lighted area Spacecraft surface reflectance; solar radiation pressure surface mass ratio Initial mass of the spacecraft The thrust vector of the orbital control engine in this system Specific impulse of rail-controlled engine Initial position of the inertial frame Initial velocity of the inertial frame ; Track dynamics module storage cycle , For positive integers; allowable error limit of inertial frame position Inertial frame velocity tolerance limit ; (3.14) Variables recorded by the orbital dynamics module include: the number of orbital times. ,in The number of positions in the inertial frame. The number of inertial frame velocities ; 3.2 Attitude dynamics module design, specifically including: (3.21) Posture kinematic equations Transition sequence, in which , and Values However, they are all different. The kinematic equations for spacecraft attitude are: in, These represent the roll angle, pitch angle, and yaw angle of the spacecraft relative to its orbital system, respectively. This represents the component of the spacecraft's instantaneous rotational speed relative to the orbital system within this system. Represents Euler angular rate and The relation matrix; where , The rotational velocity of the body relative to the inertial frame is expressed in this system. The rotational speed of the orbital frame relative to the inertial frame is expressed in the orbital frame, where the Earth's gravitational constant is... km 3 / s 2 , For the semi-major axis of the spacecraft orbit, This represents the transformation matrix from the orbital system to the home system, while Indicate Euler angles at The attitude matrix corresponding to the transposed order; (3.22) Attitude dynamics equations Flexible spacecraft dynamics has two configurable types, as follows: (i) Momentum wheel control of attitude dynamics of flexible spacecraft in, The inertia matrix of the entire star; The installation matrix of the momentum wheel assembly. ; Here is the moment of inertia matrix of the momentum wheel assembly. ,in The number of momentum wheels. For the first The moment of inertia of a momentum wheel; Let be the rotational speed vector of the momentum wheel assembly, where For the first The rotational speed of the momentum wheel; and This represents the flexible coupling coefficient matrix of the vibration of the left and right solar panels on the spacecraft's rotation. ,in The modal order; and These are the modal coordinate vectors for the left and right solar panels; and This is the diagonal matrix of the modal frequencies of the left and right solar panels; and This is a diagonal matrix of the modal damping coefficients for the left and right solar panels, with diagonal elements set to 0.

005. Indicates the wheel control attitude control torque; The driving torque vector of the momentum wheel assembly; The driving torque noise of the momentum wheel assembly, where This indicates that the mean is 0 and the variance is... Noise level, in Nm; This is the diagonal matrix in the direction of the momentum wheel assembly's rotational speed. ; This is the set value of the dynamic friction torque of the momentum wheel assembly; The damping friction coefficient of the momentum wheel is ; the nominal rotational speed vector of the momentum wheel assembly is . The operating speed vector range of the momentum wheel assembly is: ,in , ; Represents the external disturbance torque, where This represents the triaxial representation of the solar pressure torque within this system. ,and , and , This is the optical pressure torque coefficient, which is related to the geometry and mass distribution characteristics of the spacecraft. The value is the representation of the track control interference torque in this system, and is the bound value. (ii) Attitude dynamics of jet-controlled flexible spacecraft Among them, the The lever arm vector of the attitude control thruster The thrust vector of the three axes in this system , Indicates the first The jet pulse sequence of an attitude control thruster is defined as follows: The assignment logic of the attitude control thruster command in the tested GNC software is determined. Indicates the first The attitude control thruster in the first Jet pulse width per control cycle ; (3.23) Attitude dynamics module configuration parameters, including: sequence transition It includes 6 types of sequence transformations: , , , , , Inertia matrix of the entire star , Flexible coupling coefficient matrix of left and right solar panel vibrations to the rotation of the central rigid body and , Modal order Modal frequencies of the left and right solar panels and Diagonal matrix of modal damping coefficients for left and right solar panels and The diagonal elements are set to 0.005; the number of momentum wheels. Moment of inertia of the momentum wheel , ; Moment of inertia matrix of momentum wheel assembly , Installation matrix of momentum wheel assembly dimensionality Momentum wheel assembly driving torque noise Momentum wheel assembly speed measurement noise The dynamic friction torque of the momentum wheel assembly is set to [value]. The damping friction coefficient of the momentum wheel Number of attitude control thrusters Attitude control thruster nozzle installation position , Attitude control thruster nozzle vector angle , Nominal force of attitude control thruster , ;No. The specific impulse of the individual attitude control thruster , Initial attitude angle Initial attitude angular velocity Attitude dynamics module data storage cycle , For positive integers, the tolerance limit for attitude angles is... Attitude angular velocity tolerance limit ; (3.24) Variables recorded by the attitude dynamics module include: storage of three-axis attitude angles. ,in ,in ; Storage of three-axis attitude angular velocities Total number of optimization tests for the attitude dynamics module ; 3.3 Actuator module design, specifically including: (3.31) Momentum wheel design: based on Calculate The rotational speed of the momentum wheel assembly composed of momentum wheels Then we get Measuring the rotational speed of the momentum wheel The output is given to the GNC software under test; among which For momentum wheel assembly speed measurement noise, The mean is 0 and the variance is noise, ; (3.32) Attitude control thruster design: according to Calculate the attitude control torques acting on the three axes of the spacecraft's main system during the current control cycle, and then drive the spacecraft's attitude evolution; (3.33) Track control engine design: according to Calculate the orbital control forces applied to the three axes of the spacecraft's inertial frame during the current control period, and then drive the spacecraft's orbital evolution; 3.4 Sensor module design, specifically including: (3.41) Star Sensor Design: For star sensors The attitude dynamics module outputs three-axis attitude angles. Add three-axis attitude equivalent measurement noise ,get ,according to The attitude matrix of this system relative to the orbital system is obtained by transposing the order. Then according to Calculation of star sensitivity Measured spacecraft attitude matrix Convert to the corresponding quaternion Then we get Xingmin Measured spacecraft attitude quaternions ;in , and The mean is 0 and the variance is noise, Xingmin Installation matrix, number of star sensors , ; (3.42) Gyroscope design: according to Calculate the first The angle increment of the gyroscope only ;in The three-axis attitude angular velocity output by the attitude dynamics module; For the first Add constant drift to only the gyroscope; With a mean of 0 and a variance of Random drift; Indicates the first Mounting vector for only the gyroscope input axis. To control the period, the number of gyroscopes Furthermore, the configuration satisfies that at least three gyroscopes are not coplanar. ; (3.43) Sensor module configuration parameters, including: number of star sensors StarSensitive Installation Matrix , Three-axis attitude equivalent measurement noise , Number of spinning tops ; Gyroscope mounting vector ,in ; Gyroscope constant drift ,in ; Gyroscope measurement noise ,in ; 3.5 Numerical solution module design, specifically including: (3.51) Numerical solution algorithm: The fourth-order Runge-Kutta algorithm is used to solve the spacecraft attitude dynamics equation and orbital dynamics equation numerically. (3.52) Numerical solution step size These are the configuration parameters to be optimized. (3.53) Numerical solution module configuration parameters, including: numerical solution step size ; Dynamics library optimization test time length That is, optimizing the time interval of the dynamics library through testing. ,in Based on the configuration of the GNC software under test; the actual duration of the GNC software function under test. That is, the actual time interval of the tested GNC software function is ,in Based on the functional configuration of the GNC software under test; (3.54) Variables recorded by the numerical solution module include: the total number of optimization tests performed by the numerical solution module. ; 3.6 Interface design between functional modules of the spacecraft dynamics library, including: (3.61) Inputs to the orbit dynamics module include: the current state of the orbit dynamics equations and the orbit control forces; (3.62) The output of the orbital dynamics module includes: the state of the orbital dynamics equations at the previous long time step; (3.63) Inputs to the attitude dynamics module include: the current state of the attitude dynamics equations and the attitude control torque; (3.64) The output of the attitude dynamics module includes: the state of the attitude dynamics equations at the previous long time step, the spacecraft attitude angles and attitude angular velocities; (3.65) The outputs of the actuator module include: attitude control torque and track control force; (3.66) Inputs to the sensor module include: spacecraft attitude angle and attitude angular velocity; (3.67) The inputs to the numerical solution module include: the state of the attitude dynamics equations at the previous long time step and the state of the orbital dynamics equations at the previous long time step; (3.68) The output of the numerical solution module includes: the current state of the attitude dynamics equation and the current state of the orbit dynamics equation.

2. The method for constructing a rapid integrated test dynamics library for GNC software of high-orbit spacecraft according to claim 1, characterized in that: The programming implementation of the spacecraft dynamics library software includes: implementing it using a programming language consistent with the programming language of the GNC software under test, according to the designed functional modules of the spacecraft dynamics library and the interfaces between the designed functional modules of the spacecraft dynamics library; The interaction interface between the spacecraft dynamics library software and the GNC software under test includes: Orbit time output by the orbit dynamics module Spacecraft position and speed Provide the GNC software to be tested; The spacecraft attitude quaternion measured by the star sensor output from the sensor module. Angular increment measured by gyroscope Provide the GNC software to be tested; The momentum wheel output from the actuator module measures the rotational speed. Provide the GNC software to be tested; Wheel control attitude control torque input to the actuator module Attitude control jet pulse width Orbit control jet pulse width Output by the GNC software being tested.

3. The method for constructing a rapid integrated test dynamics library for GNC software of high-orbit spacecraft according to claim 1, characterized in that: The optimized configuration of the orbital dynamics module, actuator module, and sensor module of the spacecraft dynamics library includes: Orbital Dynamics Module Optimization Configuration: The orbital dynamics module is configured with Earth's central gravitational acceleration. If the tested GNC software function involves orbit control jet thrust, then the orbit dynamics module should be configured with the orbit control jet thrust item. Otherwise, do not configure. If the tested GNC software function involves longitude... Peaceful longitude change rate Then the orbital dynamics module is configured with Earth's gravitational perturbation terms. Otherwise, do not configure. If the GNC software function being tested involves tilt angle... Right ascension of ascending node Then the orbital dynamics module is configured with solar gravitational perturbation terms. and lunar gravitational perturbation terms Otherwise, do not configure. and If the tested GNC software function involves eccentricity... Perigeal argument Then the orbital dynamics module is configured with solar radiation pressure perturbation terms. Otherwise, do not configure. ; Actuator module optimization configuration: Determine actuator module configuration parameters based on the functions of the GNC software under test: number of momentum wheels Moment of inertia of the momentum wheel , ; Moment of inertia matrix of momentum wheel assembly , Installation matrix of momentum wheel assembly dimensionality Momentum wheel assembly driving torque noise Momentum wheel assembly speed measurement noise The dynamic friction torque of the momentum wheel assembly is set to [value]. The damping friction coefficient of the momentum wheel Number of attitude control thrusters Attitude control thruster nozzle installation position , Attitude control thruster nozzle vector angle , Nominal force of attitude control thruster , ;No. The specific impulse of the individual attitude control thruster , The thrust vector of the orbital control engine in this system Specific impulse of rail-controlled engine ; Sensor module optimization configuration: Determine sensor module configuration parameters based on the functions of the GNC software under test: number of star sensors. StarSensitive Installation Matrix , Three-axis attitude equivalent measurement noise , Number of spinning tops ; Gyroscope mounting vector ,in ; Gyroscope constant drift ,in ; Gyroscope measurement noise ,in .

4. The method for constructing a rapid integrated test dynamics library for GNC software of high-orbit spacecraft according to claim 1, characterized in that: The initial configuration of the attitude dynamics module and numerical solution module of the spacecraft dynamics library includes: Initial configuration of the attitude dynamics module: If the GNC software under test involves spacecraft position and velocity, but not spacecraft attitude angles and attitude angular velocities, and wheel-controlled attitude control is used, then the attitude dynamics module is configured directly. , In other words, the attitude dynamics module no longer performs numerical calculations; if the tested GNC software function involves spacecraft attitude angles and attitude angular velocities, and wheel-controlled attitude control is used without orbit control, then rigid body spacecraft attitude dynamics is configured, i.e., configured... and Wheel-controlled attitude control torque and solar pressure torque; if the measured GNC software function involves spacecraft attitude angles and attitude angular velocities, and wheel-controlled attitude control and jet orbital control are used, then rigid body spacecraft attitude dynamics should be configured, i.e., configured... and Wheel control attitude control torque, solar radiation pressure torque, and orbital control disturbance torque; if the tested GNC software function uses jet attitude control and has no orbital control, then configure the flexible spacecraft attitude dynamics and jet attitude control torque; among which, configure the modal order. External disturbance torque If the tested GNC software function uses jet attitude control and jet orbit control, then configure the flexible spacecraft attitude dynamics, jet attitude control torque, and orbit control disturbance torque; among which, configure the modal order. External disturbance torque ; Initial configuration of the numerical solution module: The numerical solution algorithm adopts the fourth-order Runge-Kutta algorithm, and the initial step size for numerical solution is configured. .

5. The method for constructing a rapid integrated test dynamics library for GNC software of high-orbit spacecraft according to claim 1, characterized in that: The currently configured spacecraft dynamics library software refers to the spacecraft dynamics library software whose orbital dynamics module is configured with optimized settings and configured with the orbital initial time (UTC). Initial position of the inertial frame Initial velocity of the inertial frame ; The actuator module is optimized and configured with momentum wheel set driving torque noise. Nm, momentum wheel assembly speed measurement noise rpm; the sensor module adopts an optimized configuration and is configured with three-axis attitude equivalent measurement noise. rad, Gyroscope measurement noise , The attitude dynamics module uses an initial configuration and configures the initial attitude angles. Initial attitude angular velocity ; The numerical solution module uses an initial configuration and optimizes the test time using the spacecraft dynamics library. ; The integrated test environment refers to the closed-loop connection between the spacecraft dynamics library software and the GNC software under test according to the interaction interface between the spacecraft dynamics library software and the GNC software under test.

6. The method for constructing a rapid integrated test dynamics library for GNC software of high-orbit spacecraft according to claim 1, characterized in that: The process of determining the optimal configuration of the spacecraft dynamics library attitude dynamics module through testing includes: If the GNC software under test uses wheel-based attitude control, the optimized configuration of the attitude dynamics module should use the initial configuration of the attitude dynamics module; test once using the integrated test environment: record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , , for and Multiples of the least common multiple of the function This indicates that the independent variable is rounded down to negative infinity; Record the total number of optimization tests for the attitude dynamics module. ; If the GNC software under test uses jet attitude control, the steps to determine the optimal configuration of the attitude dynamics module through testing in an integrated test environment are as follows: Test 1: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , , for and Multiples of the least common multiple; Second test: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , ;if , , and and , Then determine the attitude dynamics module configuration of the flexible spacecraft attitude dynamics and the minimum modal order. Otherwise, conduct a third test; Test 3: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , ;if , , and and , Then determine the attitude dynamics module configuration of the flexible spacecraft attitude dynamics and the minimum modal order. Otherwise, conduct a fourth test; Test 4: Configuration Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , ;if , , and and , Then determine the attitude dynamics module configuration of the flexible spacecraft attitude dynamics and the minimum modal order. Otherwise, determine that the attitude dynamics module is configured as flexible spacecraft attitude dynamics with the minimum modal order. ; The total number of times the attitude dynamics module optimization test was recorded was: .

7. The method for constructing a rapid integrated test dynamics library for GNC software of high-orbit spacecraft according to claim 1, characterized in that: The optimal configuration of the spacecraft dynamics library numerical solution module was determined through testing, including: Test 1: Configuration ms, recording the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , and ;like , and , and , Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, conduct a second test; Second test: Configuration (ms), where Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , and ;like , and , and , Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, proceed to the next step. This test, among which ; No. Second test: Configuration ,in Record the three-axis attitude angles in the attitude dynamics module. and attitude angular velocity Record the position of the inertial frame in the orbital dynamics module. and speed ;in , and ;if , and , and , Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, proceed to the next step. Second test; No. Second test: Configuration The attitude dynamics module records the three-axis attitude angles. and attitude angular velocity The orbital dynamics module records the position of the inertial frame. and speed ,in , and ;like , , , , , Then determine the maximum numerical solution step size of the numerical solution module. Otherwise, determine the maximum numerical solution step size of the numerical solution module. ; The total number of optimization tests recorded for the numerical solution module is .

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