A non-cooperative target satellite rendezvous orbit design method based on active disturbance rejection control

Abstract

The application discloses a non-cooperative target satellite rendezvous orbit design method based on active disturbance rejection control and belongs to the technical field of space orbit control. In order to solve the problem of modeling and simulation of relative motion between a mobile satellite and a target satellite under the condition that the parameters of the target satellite are unknown, the application calculates initial position and velocity components in the first geodetic coordinate system of the earth center according to prior orbit parameters of the mobile satellite and the target satellite; a dynamic and kinematic model of the mobile satellite and the target satellite is established; a state equation of relative motion between the mobile satellite and the target satellite is constructed by taking relative motion position, velocity and disturbance acceleration as independent variables; an extended state observer is adopted to estimate state variables; a mobile satellite orbit motion control model is established based on a proportional and differential control framework; and disturbance acceleration is compensated by adopting active disturbance rejection design to realize the rendezvous orbit design of the non-cooperative target satellite. The application does not need to accurately obtain orbit motion characteristics of the target satellite, and the method has strong adaptability.

Description

Technical Field

[0001] This invention belongs to the field of aerospace orbit control technology, specifically relating to a method for designing rendezvous orbits of non-cooperative target satellites based on active disturbance rejection control. Background Technology

[0002] Target satellite rendezvous orbit design refers to applying orbital control forces to a maneuvering satellite within a certain period of time to change its original speed and position, so as to ensure that the maneuvering satellite and the space target satellite coincide in position within a specific time.

[0003] In the rendezvous orbit design of cooperative target satellites, parameters such as the mass characteristics, aerodynamic shape characteristics and orbital motion characteristics of the maneuvering satellite and the target satellite can be accurately obtained, and precise orbit extrapolation calculations can be performed. During the rendezvous process, precise rendezvous can be achieved by simultaneously controlling the orbits of the maneuvering satellite and the target satellite and cooperating with each other.

[0004] In rendezvous with non-cooperative target satellites, the target satellite's mass characteristics, aerodynamic shape, and orbital motion parameters are unknown. Furthermore, the target satellite may undergo orbital maneuvers during the rendezvous, making it difficult to accurately model and predict its trajectory. This hinders the ability to generate orbital control strategies in advance, necessitating real-time online rendezvous orbit control based on changes in the target satellite's flight trajectory. This places high demands on the speed and accuracy of the control algorithm under unknown disturbance responses. Traditional orbital rendezvous control methods have limited adaptability when the controlled object's parameters are uncertain. Summary of the Invention

[0005] The problem this invention aims to solve is the modeling and simulation of the relative motion between a maneuvering satellite and a target satellite when the target satellite's mass characteristics, aerodynamic shape characteristics, and orbital motion characteristics are unknown. It proposes a non-cooperative target satellite rendezvous orbit design method based on active disturbance rejection control.

[0006] To achieve the above objectives, the present invention provides the following technical solution:

[0007] A method for designing rendezvous orbits for non-cooperative target satellites based on active disturbance rejection control includes the following steps:

[0008] S1. Based on the prior orbital parameters of the maneuvering satellite obtained from ground telemetry and control, calculate the initial position and velocity components of the maneuvering satellite in the geocentric first equatorial coordinate system;

[0009] S2. Based on the prior orbital parameters of the target satellite obtained from ground telemetry and control, calculate the initial position and velocity components of the target satellite in the geocentric first equatorial coordinate system;

[0010] S3. Based on the initial position and velocity of the maneuvering satellite, establish the motion equations of the maneuvering satellite in the first geocentric equatorial coordinate system, and use the fourth-order Runge-Kutta integral algorithm to recursively deduce the changes in the position and velocity of the computer maneuvering satellite over time.

[0011] S4. Based on the initial position and velocity of the target satellite, establish the motion equation of the target satellite in the first geocentric equatorial coordinate system, and use the fourth-order Runge-Kutta integral algorithm to recursively calculate the changes in the position and velocity of the target satellite over time.

[0012] S5. In the geocentric first equatorial coordinate system, with the relative position and relative velocity between the maneuvering satellite and the target satellite as state variables, establish the relative motion state equation between the maneuvering satellite and the target satellite, and calculate the disturbance acceleration of the relative motion between the maneuvering satellite and the target satellite.

[0013] S6. Remove the control acceleration actively applied by the maneuvering satellite, treat the remaining components of the relative motion acceleration between the maneuvering satellite and the target satellite as perturbation acceleration, use the extended state observer (ESO) to estimate the state variables, and identify the space perturbation acceleration values ​​such as the difference in gravitational acceleration between the maneuvering satellite and the target satellite, the aerodynamic drag of the target satellite, and solar radiation pressure.

[0014] S7. Based on the proportional and differential control framework, a motion model for orbital control of a maneuvering satellite is established by using the relative position between the maneuvering satellite and the target satellite as the proportional element and the relative velocity between the maneuvering satellite and the target satellite as the differential element.

[0015] S8. Based on the disturbance acceleration of the relative motion between the maneuvering satellite and the target satellite obtained in step S5, compensate it into the maneuvering satellite orbit control motion model obtained in step S7 to complete the rendezvous orbit control design of the non-cooperative target satellite based on active disturbance rejection control.

[0016] Furthermore, the specific implementation method of step S1 includes the following steps:

[0017] S1.1. Obtain the orbital six-axis number of the maneuvering satellite, including the semi-major axis of the maneuvering satellite. eccentricity Track inclination Perigeal argument Right ascension of ascending node Closest point ;

[0018] S1.2. Calculate the mean anomaly angle of the maneuvering satellite at time t. Angle of proximity True near point angle Earth's distance Latitude Aspect ;

[0019] S1.3. Calculate the x-axis component of the position of the maneuvering satellite at time t. y-axis components z-axis component ;

[0020] S1.4. Calculate the velocity of the maneuvering satellite at time t. x-axis component of the velocity of the maneuvering satellite at time t y-axis component y-axis component .

[0021] Furthermore, the specific implementation method of step S2 includes the following steps:

[0022] S2.1. Obtain the six orbital elements of the target satellite, including its semi-major axis. eccentricity Track inclination Perigeal argument Right ascension of ascending node Closest point ;

[0023] S2.2. Calculate the mean anomaly angle of the maneuvering satellite at time t. Angle of proximity True near point angle Earth's distance Latitude Aspect ;

[0024] S2.3. Calculate the x-axis component of the target satellite's position at time t. y-axis components z-axis component ;

[0025] S2.4. Calculate the velocity of the target satellite at time t. x-axis component of the target satellite's velocity at time t y-axis component y-axis component .

[0026] Furthermore, the specific implementation method of step S3 includes the following steps:

[0027] S3.1. Initial position and velocity of the computer-controlled satellite;

[0028] Computer-controlled satellite subjected to non-spherical perturbations of the Earth's x-axis, y-axis, and z-axis accelerations , , ; Calculate the x-axis, y-axis, and z-axis accelerations caused by atmospheric drag on a maneuvering satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by the Sun's gravity on the maneuvering satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by the Moon's gravity on the maneuvering satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by solar radiation pressure on the maneuvering satellite. , , ;

[0029] S3.2. Establish the equations of motion for the maneuvering satellite in the geocentric first equatorial coordinate system, expressed as:

[0030]

[0031] in, , , Actively applied control acceleration for maneuvering satellites is the Earth's gravitational constant.

[0032] Furthermore, the specific implementation method of step S4 includes the following steps:

[0033] S4.1. Calculate the initial position and velocity of the target satellite;

[0034] Calculate the x-axis, y-axis, and z-axis accelerations of the target satellite due to the Earth's non-spherical perturbation. , , ; Calculate the x-axis, y-axis, and z-axis accelerations caused by atmospheric drag on the target satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by the Sun's gravity on the target satellite. , , Calculate the x-axis, y-axis, and z-axis perturbation accelerations exerted by the Moon's gravity on the target satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by solar radiation pressure on the target satellite. , , ;

[0035] S4.2. Establish the equations of motion for the target satellite in the geocentric first equatorial coordinate system, expressed as:

[0036]

[0037] in, , , The acceleration applied to the target satellite's maneuver.

[0038] Furthermore, the specific implementation method of step S5 includes the following steps:

[0039] S5.1. In the geocentric first equatorial coordinate system, set the state variables of the x-axis, y-axis, and z-axis in the three directions as follows: , , The control variables for the x-axis, y-axis, and z-axis are: , , The observations for the x-axis, y-axis, and z-axis are: , , ,but:

[0040]

[0041]

[0042]

[0043] in, , , These are the projections of the relative positions of the maneuvering satellite and the target satellite in three directions within the geocentric first equatorial coordinate system; , , These are the projections of the relative velocity between the maneuvering satellite and the target satellite in three directions within the geocentric first equatorial coordinate system; , , These are the observed values ​​of the relative positions between the maneuvering satellite and the target satellite. , , These are the observed values ​​of the relative velocity between the maneuvering satellite and the target satellite; , , These are the disturbance accelerations that need to be estimated;

[0044] , , ;

[0045] S5.2. Considering the relative motion between the maneuvering satellite and the target satellite, the relative motion state equations between the maneuvering satellite and the target satellite are established, yielding:

[0046]

[0047] in, , , , , , , , , These are the coefficients of the state equation;

[0048] , , .

[0049] Furthermore, the specific implementation method of step S6 includes the following steps:

[0050] S6.1. Calculate the coefficients of the extended state observer. , , ,get:

[0051]

[0052]

[0053]

[0054] in, This indicates that the eigenvalues ​​of the matrix within the parentheses are calculated. , , This represents the bandwidth for observing the velocities of the maneuvering satellite relative to the target satellite in three directions. , , This represents the bandwidth for observing the positions of the maneuvering satellite and the target satellite in three directions relative to each other. , , This represents the bandwidth used to observe the interference acceleration in the three directions of the relative motion between the maneuvering satellite and the target satellite;

[0055] S6.2. Construct an extended state observer (ESO) to estimate the state variables through observation, then:

[0056]

[0057] in, , , They are respectively , , The estimated value.

[0058] Furthermore, the orbital control motion model for the maneuvering satellite established in step S7 is as follows:

[0059]

[0060]

[0061]

[0062] in, , , They are respectively , , The estimated value, , , They are respectively , , The estimated value;

[0063] , , , , , ;

[0064] in, , These are the position feedback gain and velocity feedback gain in the X direction of the first geocentric equatorial coordinate system, respectively. , These are the position feedback gain and velocity feedback gain in the Y direction of the first geocentric equatorial coordinate system, respectively. , These are the position feedback gain and velocity feedback gain in the Z direction of the first equatorial coordinate system at the Earth's center; they are comprehensively adjusted through the design of the control tracking transient process.

[0065] Furthermore, the expression for the non-cooperative target satellite rendezvous orbit control design based on active disturbance rejection control in step S8 is as follows:

[0066]

[0067]

[0068]

[0069] in, , , These are the estimated values ​​of the perturbation acceleration in the X, Y, and Z directions of the first geocentric equatorial coordinate system, respectively.

[0070] , , .

[0071] The beneficial effects of this invention are:

[0072] This invention discloses a non-cooperative target satellite rendezvous orbit design method based on active disturbance rejection control. It employs an extended state observer to estimate online the relative acceleration deviations caused by gravity, aerodynamic drag, solar radiation pressure, etc., resulting in relative motion between the maneuvering satellite and the target satellite, as well as the orbital acceleration actively applied by the target satellite. These deviations are then compensated for in the guidance commands using a feedforward approach. This method solves the problem of modeling and simulating the relative motion between the maneuvering and target satellites when parameters such as the target satellite's mass characteristics, aerodynamic shape characteristics, and orbital motion characteristics are unknown. It exhibits a fast response to target satellite maneuvering avoidance during rendezvous, significantly improving the robustness of the control algorithm. It can be effectively applied to space orbit control design.

[0073] The present invention provides a non-cooperative target satellite rendezvous orbit design method based on active disturbance rejection control. This method does not require accurate acquisition of the target satellite's orbital motion characteristics and can adapt to unknown maneuvers of the target satellite during the rendezvous process, thus exhibiting strong adaptability. Attached Figure Description

[0074] Figure 1 This is a flowchart of a non-cooperative target satellite rendezvous orbit design method based on active disturbance rejection control, as described in this invention.

[0075] Figure 2 This refers to the relative motion disturbance acceleration in the X direction of this invention.

[0076] Figure 3 This refers to the relative motion disturbance acceleration in the Y direction of this invention.

[0077] Figure 4 This refers to the relative motion disturbance acceleration in the Z direction of this invention.

[0078] Figure 5 The distance between the maneuvering satellite and the target satellite in this invention is denoted as . Detailed Implementation

[0079] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are only for explaining the invention and are not intended to limit the invention; that is, the described specific embodiments are merely a part of the embodiments of the invention, and not all of them. The components of the specific embodiments of the invention described and shown in the accompanying drawings can generally be arranged and designed in various different configurations, and the invention may also have other embodiments.

[0080] Therefore, the following detailed description of specific embodiments of the invention provided in the accompanying drawings is not intended to limit the scope of the claimed invention, but merely to illustrate selected specific embodiments of the invention. All other specific embodiments obtained by those skilled in the art based on these specific embodiments without inventive effort are within the scope of protection of this invention.

[0081] To further understand the invention's content, features, and effects, the following specific embodiments are provided, along with accompanying drawings. Figure 1 - Appendix Figure 5 Detailed explanation is as follows:

[0082] Example 1:

[0083] A method for designing rendezvous orbits for non-cooperative target satellites based on active disturbance rejection control includes the following steps:

[0084] S1. Based on the prior orbital parameters of the maneuvering satellite obtained from ground telemetry and control, calculate the initial position and velocity components of the maneuvering satellite in the geocentric first equatorial coordinate system;

[0085] Furthermore, the specific implementation method of step S1 includes the following steps:

[0086] S1.1. Obtain the orbital six-axis number of the maneuvering satellite, including the semi-major axis of the maneuvering satellite. eccentricity Track inclination Perigeal argument Right ascension of ascending node Closest point ;

[0087] S1.2. Calculate the mean anomaly angle of the maneuvering satellite at time t. Angle of proximity True near point angle Earth's distance Latitude Aspect ;

[0088] Furthermore, calculate the mean anomaly angle of the maneuvering satellite at time t. Then we have:

[0089] ;

[0090] in, Let be the Earth's gravitational constant, with values ​​ranging from 1 to 10. ;

[0091] Calculate the perigee angle of the maneuvering satellite at time t. Solving Kepler's equations using an iterative method yields:

[0092] ;

[0093] when ,Pick , For a given computational precision, the initial value for iteration Value ;

[0094] Calculate the true perimeter angle of the maneuvering satellite at time t. Then we have:

[0095] ;

[0096] Calculate the geocentric distance of the maneuvering satellite at time t. Then we have:

[0097] ;

[0098] Calculate the latitudinal argument of the maneuvering satellite at time t. Then we have:

[0099] ;

[0100] S1.3. Calculate the x-axis component of the position of the maneuvering satellite at time t. y-axis components z-axis component Then we have:

[0101] ;

[0102] Calculate the velocity of the maneuvering satellite at time t. x-axis component of the velocity of the maneuvering satellite at time t y-axis component y-axis component Then we have:

[0103]

[0104]

[0105]

[0106] ;

[0107] S2. Based on the prior orbital parameters of the target satellite obtained from ground telemetry and control, calculate the initial position and velocity components of the target satellite in the geocentric first equatorial coordinate system;

[0108] Furthermore, the specific implementation method of step S2 includes the following steps:

[0109] S2.1. Obtain the six orbital elements of the target satellite, including its semi-major axis. eccentricity Track inclination Perigeal argument Right ascension of ascending node Closest point ;

[0110] S2.2. Calculate the mean anomaly angle of the maneuvering satellite at time t. Angle of proximity True near point angle Earth's distance Latitude Aspect ;

[0111] Furthermore, calculate the mean anomaly angle of the target satellite at time t. Then we have:

[0112] ;

[0113] Calculate the perigee angle of the target satellite at time t. Solving Kepler's equations using an iterative method yields:

[0114]

[0115] when ,Pick , For a given computational precision, the initial value for iteration Value .

[0116] Calculate the true anomaly angle of the target satellite at time t. Then we have:

[0117]

[0118] Calculate the geocentric distance of the target satellite at time t. Then we have:

[0119]

[0120] Calculate the latitude argument of the target satellite at time t. Then we have:

[0121] ;

[0122] S2.3. Calculate the x-axis component of the target satellite's position at time t. y-axis components z-axis component Then we have:

[0123] ;

[0124] S2.4. Calculate the velocity of the target satellite at time t. x-axis component of the target satellite's velocity at time t y-axis component y-axis component Then we have:

[0125]

[0126]

[0127]

[0128] .

[0129] S3. Based on the initial position and velocity of the maneuvering satellite, establish the motion equations of the maneuvering satellite in the first geocentric equatorial coordinate system, and use the fourth-order Runge-Kutta integral algorithm to recursively deduce the changes in the position and velocity of the computer maneuvering satellite over time.

[0130] Furthermore, the specific implementation method of step S3 includes the following steps:

[0131] S3.1. Initial position and velocity of the computer-controlled satellite;

[0132] Computer-controlled satellite subjected to non-spherical perturbations of the Earth's x-axis, y-axis, and z-axis accelerations , , ; Calculate the x-axis, y-axis, and z-axis accelerations caused by atmospheric drag on a maneuvering satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by the Sun's gravity on the maneuvering satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by the Moon's gravity on the maneuvering satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by solar radiation pressure on the maneuvering satellite. , , ;

[0133] Furthermore, the computer-controlled satellite experiences non-spherical perturbation acceleration from the Earth. , , Then we have:

[0134]

[0135] in, The radius of the Earth's equator is taken as 6,378,140 m. The harmonic coefficient of Earth's gravitational field;

[0136] Value .

[0137] Calculate the acceleration caused by atmospheric drag on a maneuvering satellite , , Then we have:

[0138]

[0139]

[0140] in, The drag coefficient of the maneuvering satellite, Atmospheric density, The characteristic area of ​​a mobile satellite, For the mass of the mobile satellite, Let be the Earth's rotational angular velocity, and take the value. ;

[0141] Calculate the perturbation acceleration caused by the Sun's gravity on a maneuvering satellite , , Then we have:

[0142]

[0143]

[0144]

[0145] in, The mass ratio of the Sun to the Earth is 332946.0. The distance between the sun and the moving satellite; , , These are the coordinate components of the Sun in the first geocentric equatorial system; The distance between the Earth and the Sun;

[0146] Calculate the perturbation acceleration caused by the Moon's gravity on the maneuvering satellite , , Then we have:

[0147]

[0148]

[0149]

[0150] in, The Earth-Moon mass ratio is taken as 0.01230002; The distance between the moon and the mobile satellite; , , These are the coordinate components of the Moon in the first geocentric equatorial system; This represents the distance between the Earth and the Sun.

[0151] Calculate the perturbation acceleration caused by solar radiation pressure on a moving satellite , , Then we have:

[0152]

[0153] in, The solar radiation pressure constant has the following values: ; The surface reflection characteristic coefficient of the maneuvering satellite is 1 for total absorption, 1.44 for complete diffuse reflection, and 2 for complete specular reflection. The projected area of ​​the mobile satellite in the direction perpendicular to sunlight;

[0154] S3.2. Establish the equations of motion for the maneuvering satellite in the geocentric first equatorial coordinate system, expressed as:

[0155]

[0156] in, , , Actively applied control acceleration for maneuvering satellites is the Earth's gravitational constant.

[0157] S4. Based on the initial position and velocity of the target satellite, establish the motion equation of the target satellite in the first geocentric equatorial coordinate system, and use the fourth-order Runge-Kutta integral algorithm to recursively calculate the changes in the position and velocity of the target satellite over time.

[0158] Furthermore, the specific implementation method of step S4 includes the following steps:

[0159] S4.1. Calculate the initial position and velocity of the target satellite;

[0160] Calculate the x-axis, y-axis, and z-axis accelerations of the target satellite due to the Earth's non-spherical perturbation. , , ; Calculate the x-axis, y-axis, and z-axis accelerations caused by atmospheric drag on the target satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by the Sun's gravity on the target satellite. , , Calculate the x-axis, y-axis, and z-axis perturbation accelerations exerted by the Moon's gravity on the target satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by solar radiation pressure on the target satellite. , , ;

[0161] Furthermore, the non-spherical perturbation acceleration of the Earth acting on the target satellite is calculated. , , Then we have:

[0162] ;

[0163] Calculate the acceleration of the target satellite caused by atmospheric drag. , , Then we have:

[0164]

[0165]

[0166] in, The drag coefficient of the target satellite, Atmospheric density, The characteristic area of ​​a mobile satellite, For the mass of the mobile satellite.

[0167] Calculate the perturbation acceleration caused by the Sun's gravity on the target satellite , , Then we have:

[0168]

[0169]

[0170] in, This represents the distance between the sun and the target satellite.

[0171] Calculate the perturbation acceleration of the target satellite caused by the Moon's gravity. , , Then we have:

[0172]

[0173]

[0174] in, This represents the distance between the moon and the target satellite.

[0175] Calculate the perturbation acceleration of the target satellite caused by solar radiation pressure. , , Then we have:

[0176]

[0177] in, The surface reflection characteristic coefficient of the target satellite is 1 for total absorption, 1.44 for complete diffuse reflection, and 2 for complete specular reflection. The projected area of ​​the target satellite in the direction perpendicular to sunlight;

[0178] S4.2. Establish the equations of motion for the target satellite in the geocentric first equatorial coordinate system, expressed as:

[0179]

[0180] in, , , The acceleration applied to the target satellite's maneuver.

[0181] S5. In the geocentric first equatorial coordinate system, with the relative position and relative velocity between the maneuvering satellite and the target satellite as state variables, establish the relative motion state equation between the maneuvering satellite and the target satellite, and calculate the disturbance acceleration of the relative motion between the maneuvering satellite and the target satellite.

[0182] Furthermore, the specific implementation method of step S5 includes the following steps:

[0183] S5.1. In the geocentric first equatorial coordinate system, set the state variables of the x-axis, y-axis, and z-axis in the three directions as follows: , , The control variables for the x-axis, y-axis, and z-axis are: , , The observations for the x-axis, y-axis, and z-axis are: , , ,but:

[0184]

[0185]

[0186]

[0187] in, , , These are the projections of the relative positions of the maneuvering satellite and the target satellite in three directions within the geocentric first equatorial coordinate system; , , These are the projections of the relative velocity between the maneuvering satellite and the target satellite in three directions within the geocentric first equatorial coordinate system; , , These are the observed values ​​of the relative positions between the maneuvering satellite and the target satellite. , , These are the observed values ​​of the relative velocity between the maneuvering satellite and the target satellite; , , These are the disturbance accelerations that need to be estimated;

[0188] , , ;

[0189] S5.2. Considering the relative motion between the maneuvering satellite and the target satellite, the relative motion state equations between the maneuvering satellite and the target satellite are established, yielding:

[0190]

[0191] in, , , , , , , , , These are the coefficients of the state equation;

[0192] , , .

[0193] S6. Remove the control acceleration actively applied by the maneuvering satellite, treat the remaining components of the relative motion acceleration between the maneuvering satellite and the target satellite as perturbation acceleration, use the extended state observer (ESO) to estimate the state variables, and identify the space perturbation acceleration values ​​such as the difference in gravitational acceleration between the maneuvering satellite and the target satellite, the aerodynamic drag of the target satellite, and solar radiation pressure.

[0194] Furthermore, the specific implementation method of step S6 includes the following steps:

[0195] S6.1. Calculate the coefficients of the extended state observer. , , ,get:

[0196]

[0197]

[0198]

[0199] in, This indicates that the eigenvalues ​​of the matrix within the parentheses are calculated. , , This represents the bandwidth for observing the velocities of the maneuvering satellite relative to the target satellite in three directions. , , This represents the bandwidth for observing the positions of the maneuvering satellite and the target satellite in three directions relative to each other. , , This represents the bandwidth used to observe the interference acceleration in the three directions of the relative motion between the maneuvering satellite and the target satellite;

[0200] S6.2. Construct an extended state observer (ESO) to estimate the state variables through observation, then:

[0201]

[0202] in, , , They are respectively , , The estimated value.

[0203] S7. Based on the proportional and differential control framework, a motion model for orbital control of a maneuvering satellite is established by using the relative position between the maneuvering satellite and the target satellite as the proportional element and the relative velocity between the maneuvering satellite and the target satellite as the differential element.

[0204] Furthermore, the orbital control motion model for the maneuvering satellite established in step S7 is as follows:

[0205]

[0206]

[0207]

[0208] in, , , They are respectively , , The estimated value, , , They are respectively , , The estimated value;

[0209] , , , , , ;

[0210] in, , These are the position feedback gain and velocity feedback gain in the X direction of the first geocentric equatorial coordinate system, respectively. , These are the position feedback gain and velocity feedback gain in the Y direction of the first geocentric equatorial coordinate system, respectively. , These are the position feedback gain and velocity feedback gain in the Z direction of the first equatorial coordinate system at the Earth's center; they are comprehensively adjusted through the design of the control tracking transient process.

[0211] S8. Based on the disturbance acceleration of the relative motion between the maneuvering satellite and the target satellite obtained in step S5, compensate it into the maneuvering satellite orbit control motion model obtained in step S7 to complete the rendezvous orbit control design of the non-cooperative target satellite based on active disturbance rejection control.

[0212] Furthermore, the expression for the non-cooperative target satellite rendezvous orbit control design based on active disturbance rejection control in step S8 is as follows:

[0213]

[0214]

[0215]

[0216] in, , , These are the estimated values ​​of the perturbation acceleration in the X, Y, and Z directions of the first geocentric equatorial coordinate system, respectively.

[0217] , , .

[0218] The simulation experiment was conducted using the method of this embodiment, and the orbital parameters of the maneuvering satellite and the target satellite were set as shown in Table 1:

[0219] Table 1 Initial orbital parameters

[0220]

[0221] The mass, thrust, and aerodynamic parameters of the maneuvering satellite are set as follows: Initial mass: 100 kg; Frontal area: 2 m² 2 Radiation area: 2m 2 Three-directional thrust: 1000N;

[0222] The target satellite's mass, thrust, and aerodynamic parameters are set as follows: Mass: 5000 kg; Frontal area: 20 m² 2 Radiation area: 20m 2 Mobility: 0.1g0;

[0223] The simulation results are as follows: The initial position difference between the maneuvering satellite and the target satellite is 140 km, and the initial velocity difference is 65 m / s. At 50 s, the target satellite begins to maneuver in three directions, with accelerations of -1 m / s² in the X, Y, and Z directions. 2 0.5 m / s 2 -0.8 m / s 2 By employing the above-mentioned orbit control method, the maneuvering satellite completed its orbital rendezvous with the target satellite in 282 seconds, with a relative position error of less than 0.1m and a relative velocity error of less than 0.01m / s.

[0224] It should be noted that relational terms such as "first" and "second" are used merely to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0225] Although this application has been described above with reference to specific embodiments, various modifications can be made and components can be replaced with equivalents without departing from the scope of this application. In particular, as long as there is no structural conflict, the features in the specific embodiments disclosed in this application can be combined with each other in any way. The lack of an exhaustive description of these combinations in this specification is merely for the sake of brevity and resource conservation. Therefore, this application is not limited to the specific embodiments disclosed herein, but includes all technical solutions falling within the scope of the claims.

Claims

1. A method for designing rendezvous orbits for non-cooperative target satellites based on active disturbance rejection control, characterized in that, Includes the following steps: S1. Based on the prior orbital parameters of the maneuvering satellite obtained from ground telemetry and control, calculate the initial position and velocity components of the maneuvering satellite in the geocentric first equatorial coordinate system; S2. Based on the prior orbital parameters of the target satellite obtained from ground telemetry and control, calculate the initial position and velocity components of the target satellite in the geocentric first equatorial coordinate system; S3. Based on the initial position and velocity of the maneuvering satellite, establish the motion equations of the maneuvering satellite in the first geocentric equatorial coordinate system, and use the fourth-order Runge-Kutta integral algorithm to recursively deduce the changes in the position and velocity of the computer maneuvering satellite over time. S4. Based on the initial position and velocity of the target satellite, establish the motion equation of the target satellite in the first geocentric equatorial coordinate system, and use the fourth-order Runge-Kutta integral algorithm to recursively calculate the changes in the position and velocity of the target satellite over time. S5. In the geocentric first equatorial coordinate system, with the relative position and relative velocity between the maneuvering satellite and the target satellite as state variables, establish the relative motion state equation between the maneuvering satellite and the target satellite, and calculate the disturbance acceleration of the relative motion between the maneuvering satellite and the target satellite. S6. Remove the control acceleration actively applied by the maneuvering satellite, treat the remaining components of the relative motion acceleration between the maneuvering satellite and the target satellite as perturbation acceleration, use the extended state observer (ESO) to estimate the state variables, and identify the differences in gravitational acceleration between the maneuvering satellite and the target satellite, the aerodynamic drag of the target satellite, and the values ​​of solar radiation pressure space perturbation acceleration. S7. Based on the proportional and differential control framework, a motion model for orbital control of a maneuvering satellite is established by using the relative position between the maneuvering satellite and the target satellite as the proportional element and the relative velocity between the maneuvering satellite and the target satellite as the differential element. S8. Based on the disturbance acceleration of the relative motion between the maneuvering satellite and the target satellite obtained in step S5, compensate it into the maneuvering satellite orbit control motion model obtained in step S7 to complete the rendezvous orbit control design of the non-cooperative target satellite based on active disturbance rejection control.

2. The method for designing rendezvous orbits of non-cooperative target satellites based on active disturbance rejection control according to claim 1, characterized in that, The specific implementation method of step S1 includes the following steps: S1.

1. Obtain the orbital six-axis number of the maneuvering satellite, including the semi-major axis of the maneuvering satellite. eccentricity Track inclination Perigeal argument Right ascension of ascending node Closest point ; S1.

2. Calculate the mean anomaly angle of the maneuvering satellite at time t. Angle of proximity True near point angle Earth's distance Latitude Aspect ; S1.

3. Calculate the x-axis component of the position of the maneuvering satellite at time t. y-axis components z-axis component ; S1.

4. Calculate the velocity of the maneuvering satellite at time t. x-axis component of the velocity of the maneuvering satellite at time t y-axis component y-axis component .

3. The method for designing rendezvous orbits of non-cooperative target satellites based on active disturbance rejection control according to claim 2, characterized in that, The specific implementation method of step S2 includes the following steps: S2.

1. Obtain the six orbital elements of the target satellite, including its semi-major axis. eccentricity Track inclination Perigeal argument Right ascension of ascending node Closest point ; S2.

2. Calculate the mean anomaly angle of the maneuvering satellite at time t. Angle of proximity True near point angle Earth's distance Latitude Aspect ; S2.

3. Calculate the x-axis component of the target satellite's position at time t. y-axis components z-axis component ; S2.

4. Calculate the velocity of the target satellite at time t. x-axis component of the target satellite's velocity at time t y-axis component y-axis component .

4. The method for designing rendezvous orbits of non-cooperative target satellites based on active disturbance rejection control according to claim 3, characterized in that, The specific implementation method of step S3 includes the following steps: S3.

1. Initial position and velocity of the computer-controlled satellite; Computer-controlled satellite subjected to non-spherical perturbations of the Earth's x-axis, y-axis, and z-axis accelerations , , ; Calculate the x-axis, y-axis, and z-axis accelerations caused by atmospheric drag on a maneuvering satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by the Sun's gravity on the maneuvering satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by the Moon's gravity on the maneuvering satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by solar radiation pressure on the maneuvering satellite. , , ; S3.

2. Establish the equations of motion for the maneuvering satellite in the geocentric first equatorial coordinate system.

5. The method for designing rendezvous orbits of non-cooperative target satellites based on active disturbance rejection control according to claim 4, characterized in that, The specific implementation method of step S4 includes the following steps: S4.

1. Calculate the initial position and velocity of the target satellite; Calculate the x-axis, y-axis, and z-axis accelerations of the target satellite due to the Earth's non-spherical perturbation. , , ; Calculate the x-axis, y-axis, and z-axis accelerations caused by atmospheric drag on the target satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by the Sun's gravity on the target satellite. , , Calculate the x-axis, y-axis, and z-axis perturbation accelerations exerted by the Moon's gravity on the target satellite. , , ; Calculate the x-axis, y-axis, and z-axis perturbation accelerations caused by solar radiation pressure on the target satellite. , , ; S4.

2. Establish the equations of motion for the target satellite in the geocentric first equatorial coordinate system.

6. The method for designing rendezvous orbits of non-cooperative target satellites based on active disturbance rejection control according to claim 5, characterized in that, The specific implementation method of step S5 includes the following steps: S5.

1. In the geocentric first equatorial coordinate system, set the state variables of the x-axis, y-axis, and z-axis in the three directions as follows: , , The control variables for the x-axis, y-axis, and z-axis are: , , The observations for the x-axis, y-axis, and z-axis are: , , ,but: ； ； ； in, , , These are the projections of the relative positions of the maneuvering satellite and the target satellite in three directions within the geocentric first equatorial coordinate system; , , These are the projections of the relative velocity between the maneuvering satellite and the target satellite in three directions within the geocentric first equatorial coordinate system; , , These are the observed values ​​of the relative positions between the maneuvering satellite and the target satellite. , , These are the observed values ​​of the relative velocity between the maneuvering satellite and the target satellite; , , These are the disturbance accelerations that need to be estimated; ， ， ； S5.

2. Considering the relative motion between the maneuvering satellite and the target satellite, the relative motion state equations between the maneuvering satellite and the target satellite are established, yielding: ； in, , , , , , , , , These are the coefficients of the state equation.

7. The method for designing rendezvous orbits of non-cooperative target satellites based on active disturbance rejection control according to claim 6, characterized in that, The specific implementation method of step S6 includes the following steps: S6.

1. Calculate the coefficients of the extended state observer. , , ,get: ； ； ； in, This indicates that the eigenvalues ​​of the matrix within the parentheses are calculated. , , This represents the bandwidth for observing the velocities of the maneuvering satellite relative to the target satellite in three directions. , , This represents the bandwidth for observing the positions of the maneuvering satellite and the target satellite in three directions relative to each other. , , This represents the bandwidth used to observe the interference acceleration in the three directions of the relative motion between the maneuvering satellite and the target satellite; S6.

2. Construct an extended state observer (ESO) to estimate the state variables through observation, then: ； in, , , They are respectively , , The estimated value.

8. The method for designing rendezvous orbits of non-cooperative target satellites based on active disturbance rejection control according to claim 7, characterized in that, The orbital control motion model of the maneuvering satellite established in step S7 is as follows: ； ； ； in, , , They are respectively , , The estimated value, , , They are respectively , , The estimated value; ， ， ， ， ， ； in, , These are the position feedback gain and velocity feedback gain in the X direction of the first geocentric equatorial coordinate system, respectively. , These are the position feedback gain and velocity feedback gain in the Y direction of the first geocentric equatorial coordinate system, respectively. , These are the position feedback gain and velocity feedback gain in the Z direction of the first equatorial coordinate system at the Earth's center; they are comprehensively adjusted through the design of the control tracking transient process.

9. The method for designing rendezvous orbits of non-cooperative target satellites based on active disturbance rejection control according to claim 8, characterized in that, The expression for the non-cooperative target satellite rendezvous orbit control design based on active disturbance rejection control in step S8 is: ； ； ； in, , , These are the estimated values ​​of the perturbation acceleration in the X, Y, and Z directions of the first geocentric equatorial coordinate system, respectively. ， ， 。

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