阅读说明：本技术 一种微纳卫星在轨自主交会控制的方法及计算机设备 (Method for controlling in-orbit autonomous intersection of micro/nano satellite and computer equipment ) 是由 徐�明 白雪 郑亚茹 胡海霞 严晗 杨志 于灵惠 马林 于 2021-08-02 设计创作，主要内容包括：本申请公开了一种微纳卫星在轨自主交会控制的方法及计算机设备,该方法包括：分别根据相对动力学模型和绝对动力学模型计算得到相对导航数据和绝对导航数据,根据相对导航数据以及绝对导航数据计算相对平均轨道根数；根据相对导航数据、绝对导航数据以及相对平均轨道根数进行预设时间段内的轨道递推得到轨道递推数据,根据轨道递推数据判断是否达到预设交会条件；若达到,则计算追踪航天器在初始时刻和终点时刻的两次控制脉冲,根据轨控指令集以及两次控制脉冲对追踪航天器和目标航天器的相对轨道参数进行调整,以使得追踪航天器和目标航天器进行自主交会。本申请解决了现有技术中航天器变轨交会对接控制精度较低的技术问题。(The application discloses a method and computer equipment for on-orbit autonomous rendezvous control of a micro-nano satellite, wherein the method comprises the following steps: calculating to obtain relative navigation data and absolute navigation data according to the relative dynamics model and the absolute dynamics model respectively, and calculating the relative average orbit number according to the relative navigation data and the absolute navigation data; performing track recursion within a preset time period according to the relative navigation data, the absolute navigation data and the relative average track number to obtain track recursion data, and judging whether a preset rendezvous condition is met according to the track recursion data; if the target spacecraft and the tracking spacecraft meet the requirement, calculating two control pulses of the tracking spacecraft at the initial moment and the terminal moment, and adjusting relative orbit parameters of the tracking spacecraft and the target spacecraft according to the orbit control instruction set and the two control pulses so as to enable the tracking spacecraft and the target spacecraft to conduct autonomous rendezvous. The application solves the technical problem that the spacecraft orbital transfer intersection butt joint control precision is low in the prior art.)

1. A method for controlling in-orbit autonomous rendezvous of a micro-nano satellite is characterized by comprising the following steps:

the method comprises the steps of constructing an absolute dynamic model and a relative dynamic model of autonomous interaction, calculating and outputting relative navigation data according to the relative dynamic model, calculating to obtain absolute navigation data according to the absolute dynamic model, and calculating a relative average orbit root according to the relative navigation data and the absolute navigation data, wherein the relative average orbit root is the difference of the average orbit roots between a tracking spacecraft and a target spacecraft;

performing track recursion within a preset time period according to the relative navigation data, the absolute navigation data and the relative average track number to obtain track recursion data; judging whether a preset rendezvous condition is met or not according to the track recursive data;

if the target spacecraft and the tracking spacecraft meet the requirement, generating an orbit control instruction set, calculating two control pulses of the tracking spacecraft at the initial time and the terminal time, and adjusting relative orbit parameters of the tracking spacecraft and the target spacecraft according to the orbit control instruction set and the two control pulses so as to enable the tracking spacecraft and the target spacecraft to conduct autonomous rendezvous.

2. The method of claim 1, wherein calculating a relative average number of tracks from the relative navigation data and the absolute navigation data comprises:

calculating the position and the speed of the tracked spacecraft according to absolute navigation data, calculating the number of the close orbits of the tracked spacecraft according to the position and the speed, and calculating the average number of the orbits of the tracked spacecraft according to the number of the close orbits;

and obtaining a relative average orbit root by adopting Kalman filtering UKF estimation according to the relative navigation data, the close orbit root and the average orbit root.

3. The method of claim 2, wherein calculating absolute navigation data from the absolute dynamical model comprises:

determining the sum of the pulse speed after orbit control and the speed before orbit control, calculating the state quantity after orbit control according to the speed sum, updating the preset initial relative navigation data according to the state quantity, and outputting the updated relative navigation data; or

And controlling the kinematics model not to output relative navigation data within a preset time long range after orbit control until the UKF filter is ensured to reach a preset convergence effect, and outputting the relative navigation data.

4. The method of claim 3, wherein calculating two control pulses for the tracked spacecraft at the initial and end times comprises:

determining the difference of the initial semi-major axis between the tracking spacecraft and the target spacecraft, and calculating the difference of the expected semi-major axis according to the preset expected semi-major axis drift amount of each period; calculating to obtain a semi-major axis value which needs to be adjusted for tracking the spacecraft according to the difference between the initial semi-major axis and the expected semi-major axis;

calculating to obtain a pulse value required to be adjusted by the tracking spacecraft according to the semi-long axis value and a preset perturbation equation, and determining a relative motion state transfer matrix according to the semi-long axis and the pulse value;

calculating according to the relative motion state transition matrix, a preset relative motion model and positions of a preset initial moment and a preset terminal moment to obtain an initial speed and a terminal speed;

and calculating to obtain two control pulses of the tracked spacecraft at the initial moment and the end moment according to the preset initial speed and the preset end speed as well as the initial speed and the end speed.

5. The method of claim 4, wherein the tracking command set comprises: a track control instruction 1, a track control instruction 2, a track control instruction 3, a track control instruction 4, a track control instruction 5 and a track control instruction 6; wherein the content of the first and second substances,

the orbit control instruction 1 is used for controlling and changing the relative half shaft difference value into a positive value;

the orbit control instruction 2 is used for controlling and changing the relative half shaft difference value to be a negative value;

the orbit control instruction 3 and the orbit control instruction 4 are respectively used for controlling and changing the orbit control relative eccentricity vectors at the initial time and the terminal time;

the track control command 5 is used for controlling and changing the right ascension difference value of the relative ascending intersection point;

and the track control command 6 is used for controlling and changing the relative inclination difference.

6. The method of claim 5, wherein the track control command 1, the track control command 2, the track control command 3, the track control command 4, the track control command 5, and the track control command 6 each comprise: instruction word, execution time and boot time.

7. The method of claim 6, wherein the tracking command set comprises:

the tracking command set is represented by:

CIS＝[1，t₁，Δt₁；2，t₂，Δt₂；3，t₃，Δt₃；4，t₄，Δt₄；5，t₅，Δt₅；6，t₆，Δt₆；k，t_k，Δt_k]

wherein, 1, 2, 3, 4, 5 and 6 respectively represent instruction words of the track control instruction 1 to the track control instruction 6; t is t₁、t₂、t₃、t₄、t₅、t₆Respectively showing the execution time of the track control instruction 1 to the track control instruction 6; Δ t₁、Δt₂、Δt₃、Δt₄、Δt₅、Δt₆Respectively representing the startup duration corresponding to the track control instruction 1-the track control instruction 6; k represents an instruction word corresponding to the minimum execution time in the track control instructions 1 to 6; t is t_kRepresenting the minimum execution time in the track control commands 1 to 6; Δ t_kAnd the starting time corresponding to the minimum execution time instruction in the track control instructions 1 to 6 is shown.

8. The method of claim 7, wherein adjusting the relative orbit parameters of the tracking spacecraft and the target spacecraft based on the set of orbiting commands and the two control pulses comprises:

calculating a preset time delay extrapolation track parameter according to an input current track measurement parameter, and calculating to obtain track recursion data corresponding to a moment of calling an orbit control instruction set according to the extrapolation track parameter, wherein the preset time delay refers to the time delay from receiving navigation data to calling the orbit control instruction set;

and calling the orbit control instruction set according to the orbit recursion data, adjusting relative orbit parameters of the tracking spacecraft and the target spacecraft according to the orbit control instruction set and two control pulses until any orbit control instruction reaches a preset boundary threshold value, and updating the orbit control instruction set.

9. A computer device, characterized in that the computer device comprises:

a memory for storing instructions for execution by at least one processor;

a processor for executing instructions stored in a memory to perform the method of any of claims 1 to 8.

Technical Field

The application relates to the technical field of spacecraft orbit control simulation, in particular to a method and computer equipment for in-orbit autonomous rendezvous control of a micro-nano satellite.

Background

The rendezvous and docking of the spacecraft is a space strategic technology developed along with the trend of the first space technology in the 60 th century, is used as a core construction technology for the strategic operation of space civilian use and military, and is an important mark for the transition of the theoretical research success of manned aerospace to the practical engineering application. The spacecraft rendezvous and docking technology not only represents the development level of the national aerospace technology, but also is the comprehensive demonstration of the technical strength of the national aerospace technology. With the continuous progress and development of the aerospace science and technology level of the contemporary society, the rendezvous and docking technology of the spacecraft has become the mainstream and leading-edge subject of the international aerospace field, and all military and strong countries around the world have the rendezvous and docking technology of the spacecraft and the establishment of a permanent space station as the important development targets.

The spacecraft rendezvous and docking means that two spacecrafts (a tracking spacecraft and a target spacecraft) are merged on a given space orbit and are structurally connected into a whole through a special docking device. The method comprises a plurality of key contents such as orbit control, trajectory optimization, terminal guidance and the like, is a very complex space technology, and has higher requirements on reliability and precision. The target spacecraft is in a passive state without any maneuver or with little maneuver in the rendezvous and docking process, so the target spacecraft is called as a passive spacecraft; in the process, the tracking spacecraft needs to gradually approach the target spacecraft through orbital maneuver to complete docking, so the tracking spacecraft is called as an active spacecraft.

Spacecraft rendezvous and docking typically involves two phases, rendezvous and docking. The rendezvous stage is divided into a long-distance orbit maneuvering section, a short-distance guiding section and a relative position keeping (flying around) section according to the distance between the tracking spacecraft and the target spacecraft, so that the rendezvous stage belongs to the category of orbit control; the docking stage is that the tracking spacecraft starts to enter a relative position keeping (flying around) mode until the two spacecraft docking mechanisms start to contact and enable the position, the speed and the attitude to meet the docking conditions, and therefore the method belongs to the category of attitude control. The autonomous rendezvous and docking technology of the spacecraft belongs to a multidisciplinary cross topic, and is closely related to the disciplines of mathematics, mechanics, control, computers, communication engineering and the like; meanwhile, the smooth completion of autonomous rendezvous and docking of the spacecraft has high requirements on navigation, guidance control precision, system reliability and state constraint.

At present, in the conventional autonomous rendezvous and docking of the spacecraft, the relative motion relation is usually expressed by adopting an orbit number difference, the relative motion control is carried out by controlling the orbit number of the spacecraft, the long-term control effect is mainly considered, but the rendezvous and docking task time is short, so that the analysis is more visual and effective by taking the relative motion state between the spacecraft as a control target. The existing autonomous rendezvous method has the problems that the control method based on the pulse thrust action mostly adopts an open-loop control mode, the track maneuver is preset on the ground in advance, track data usually comes from absolute navigation and is easily influenced by the track maneuver, and equipment cost in the maneuver process is rarely considered.

Disclosure of Invention

The technical problem that this application was solved is: aiming at the problem that the spacecraft orbital transfer intersection butt joint control precision is low in the prior art, the application provides a method and computer equipment for in-orbit autonomous intersection control of a micro-nano satellite, in the scheme provided by the embodiment of the application, relative navigation data are calculated and output according to a relative dynamics model, absolute navigation data are calculated and obtained according to the absolute dynamics model, and a relative average orbit root is calculated according to the relative navigation data and the absolute navigation data; in other words, in the scheme provided by the embodiment of the application, absolute and relative dynamics models are respectively adopted, so that the spacecraft orbit change intersection butt joint control precision under the actual shooting condition is improved.

In a first aspect, an embodiment of the present application provides a method for controlling in-orbit autonomous rendezvous of a micro/nano satellite, where the method includes:

the method comprises the steps of constructing an absolute dynamic model and a relative dynamic model of autonomous interaction, calculating and outputting relative navigation data according to the relative dynamic model, calculating to obtain absolute navigation data according to the absolute dynamic model, and calculating a relative average orbit root according to the relative navigation data and the absolute navigation data, wherein the relative average orbit root is the difference of the average orbit roots between a tracking spacecraft and a target spacecraft;

performing track recursion within a preset time period according to the relative navigation data, the absolute navigation data and the relative average track number to obtain track recursion data, and judging whether a preset rendezvous condition is met according to the track recursion data;

if the target spacecraft and the tracking spacecraft meet the requirement, generating an orbit control instruction set, calculating two control pulses of the tracking spacecraft at the initial time and the terminal time, and adjusting relative orbit parameters of the tracking spacecraft and the target spacecraft according to the orbit control instruction set and the two control pulses so as to enable the tracking spacecraft and the target spacecraft to conduct autonomous rendezvous.

Optionally, calculating a relative average track number according to the relative navigation data and the absolute navigation data includes: calculating the position and the speed of the tracked spacecraft according to absolute navigation data, calculating the number of the close orbits of the tracked spacecraft according to the position and the speed, and calculating the average number of the orbits of the tracked spacecraft according to the number of the close orbits; and obtaining a relative average orbit root by adopting Kalman filtering UKF estimation according to the relative navigation data, the close orbit root and the average orbit root.

Optionally, calculating absolute navigation data according to the absolute dynamical model includes: determining the sum of the pulse speed after orbit control and the speed before orbit control, calculating the state quantity after orbit control according to the speed sum, updating the preset initial relative navigation data according to the state quantity, and outputting the updated relative navigation data; or controlling the kinematics model not to output relative navigation data within a preset time long range after orbit control until the UKF filter is ensured to reach a preset convergence effect, and outputting the relative navigation data.

Optionally, calculating two control pulses of the tracking spacecraft at the initial time and the end time includes: determining the difference of the initial semi-major axis between the tracking spacecraft and the target spacecraft, and calculating the difference of the expected semi-major axis according to the preset expected semi-major axis drift amount of each period; calculating to obtain a semi-major axis value which needs to be adjusted for tracking the spacecraft according to the difference between the initial semi-major axis and the expected semi-major axis; calculating to obtain a pulse value required to be adjusted by the tracking spacecraft according to the semi-long axis value and a preset perturbation equation, and determining a relative motion state transfer matrix according to the semi-long axis and the pulse value; calculating according to the relative motion state transition matrix, a preset relative motion model and positions of a preset initial moment and a preset terminal moment to obtain an initial speed and a terminal speed; and calculating to obtain two control pulses of the tracked spacecraft at the initial moment and the end moment according to the preset initial speed and the preset end speed as well as the initial speed and the end speed.

Optionally, the tracking command set comprises: a track control instruction 1, a track control instruction 2, a track control instruction 3, a track control instruction 4, a track control instruction 5 and a track control instruction 6; the track control instruction 1 is used for controlling and changing the difference value of the relative half shafts to be a positive value; the orbit control instruction 2 is used for controlling and changing the relative half shaft difference value to be a negative value; the orbit control instruction 3 and the orbit control instruction 4 are respectively used for controlling and changing the orbit control relative eccentricity vectors at the initial time and the terminal time; the track control command 5 is used for controlling and changing the right ascension difference value of the relative ascending intersection point; and the track control command 6 is used for controlling and changing the relative inclination difference.

Optionally, the track control instruction 1, the track control instruction 2, the track control instruction 3, the track control instruction 4, the track control instruction 5, and the track control instruction 6 all include: instruction word, execution time and boot time.

Optionally, the tracking command set comprises: the tracking command set is represented by:

CIS＝[1，t₁，Δt₁；2，t₂，Δt₂；3，t₃，Δt₃；4，t₄，Δt₄；5，t₅，Δt₅；6，t₆，Δt₆；k，t_k，Δt_k]

wherein, 1, 2, 3, 4, 5 and 6 respectively represent instruction words of the track control instruction 1 to the track control instruction 6; t is t₁、t₂、t₃、t₄、t₅、t₆Respectively showing the execution time of the track control instruction 1 to the track control instruction 6; Δ t₁、Δt₂、Δt₃、Δt₄、Δt₅、Δt₆Respectively representing the startup duration corresponding to the track control instruction 1-the track control instruction 6; k represents an instruction word corresponding to the minimum execution time in the track control instructions 1 to 6; t is t_kRepresenting the minimum execution time in the track control commands 1 to 6; Δ t_kAnd the starting time corresponding to the minimum execution time instruction in the track control instructions 1 to 6 is shown.

Optionally, adjusting the relative orbit parameters of the tracking spacecraft and the target spacecraft according to the orbit control instruction set and the two control pulses includes: calculating a preset time delay extrapolation track parameter according to an input current track measurement parameter, and calculating to obtain track recursion data corresponding to a moment of calling an orbit control instruction set according to the extrapolation track parameter, wherein the preset time delay refers to the time delay from receiving navigation data to calling the orbit control instruction set; and calling the orbit control instruction set according to the orbit recursion data, adjusting relative orbit parameters of the tracking spacecraft and the target spacecraft according to the orbit control instruction set and two control pulses until any orbit control instruction reaches a preset boundary threshold value, and updating the orbit control instruction set.

In a second aspect, the present application provides a computer device comprising:

a memory for storing instructions for execution by at least one processor;

a processor for executing instructions stored in a memory to perform the method of the first aspect.

Compared with the prior art, the scheme provided by the embodiment of the application has at least the following beneficial effects:

1. in the scheme provided by the embodiment of the application, relative navigation data are calculated and output according to the relative dynamics model, absolute navigation data are calculated and obtained according to the absolute dynamics model, and relative average orbit number is calculated according to the relative navigation data and the absolute navigation data; in other words, in the scheme provided by the embodiment of the application, absolute and relative dynamics models are respectively adopted, so that the spacecraft orbit change intersection butt joint control precision under the actual shooting condition is improved.

2. In the scheme provided by the embodiment of the application, relative orbit number in a nonsingular form is estimated by adopting UKF, linearization and calculation of Jacobian matrix are avoided, relative position delta r and relative speed delta v are used as random variables, absolute position r and absolute speed v are used as determination variables, and long-term maintainability of the micro-nano cluster is improved through the relative navigation design of minimum equipment cost.

Drawings

Fig. 1 is a schematic diagram of a system for in-orbit autonomous rendezvous control of a micro/nano satellite according to an embodiment of the present disclosure;

fig. 2 is a schematic flowchart of a method for controlling in-orbit autonomous rendezvous of a micro/nano satellite according to an embodiment of the present application;

FIG. 3 is a schematic diagram illustrating a process for calculating a relative average track number according to an embodiment of the present disclosure;

FIG. 4 is a schematic flow chart illustrating a process of outputting relative navigation data by a relative kinematics model according to an embodiment of the present disclosure;

fig. 5 is a schematic diagram of the distance between the tracking spacecraft star and the target spacecraft provided in the embodiment of the present application;

fig. 6 is a schematic structural diagram of a computer device according to an embodiment of the present application.

Detailed Description

In the solutions provided in the embodiments of the present application, the described embodiments are only a part of the embodiments of the present application, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application.

In order to better understand the technical solutions, the technical solutions of the present application are described in detail below with reference to the drawings and specific embodiments, and it should be understood that the specific features in the embodiments and examples of the present application are detailed descriptions of the technical solutions of the present application, and are not limitations of the technical solutions of the present application, and the technical features in the embodiments and examples of the present application may be combined with each other without conflict.

Referring to fig. 1, a schematic diagram of a system for in-orbit autonomous rendezvous control of a micro/nano satellite according to an embodiment of the present application is provided. The system comprises a target spacecraft and a tracking spacecraft, wherein in the interaction process, the tracking spacecraft tracks the target spacecraft through orbit change to realize intersection. In the system, a dynamic model can be constructed through an on-board computer on the tracking spacecraft to calculate and output relative navigation data and absolute navigation data, and the tracking spacecraft orbit control and other operations are controlled. The simulation computer can be independently arranged to construct a dynamic model to calculate and output relative navigation data and absolute navigation data, and the operations of tracking spacecraft orbit control and the like can be controlled, which is not limited herein.

The method for controlling the in-orbit autonomous rendezvous of the micro/nano satellite provided by the embodiment of the application is further described in detail with reference to the attached drawings of the specification, and the specific implementation manner of the method can include the following steps (the flow of the method is shown in fig. 2):

step 201, an absolute dynamic model and a relative dynamic model of autonomous interaction are built, relative navigation data are calculated and output according to the relative dynamic model, absolute navigation data are calculated and obtained according to the absolute dynamic model, and a relative average orbit root is calculated according to the relative navigation data and the absolute navigation data, wherein the relative average orbit root is the difference between the average orbit roots between a tracking spacecraft and a target spacecraft.

Specifically, in the solution provided in the embodiment of the present application, the dynamic model includes an absolute dynamic model and a relative dynamic model, wherein the absolute dynamic model is used to calculate the absolute navigation parameters, and the relative dynamic model is used to calculate the relative dynamic model. For example, the absolute kinetic model is a high-precision absolute kinetic model, and the precision of the absolute kinetic model can be set according to actual conditions.

Further, for a high precision absolute dynamics model, the earth's gravitational potential field can be written as follows:

in the formula:

in the formula, r is a position vector of the spacecraft under a geocentric fixed coordinate system; r_eIs the radius of the earth; g is a universal gravitation parameter; m_eIs the earth mass;and λ is the geocentric latitude and geocentric longitude of the spacecraft, respectively; p_nmIs an n-th order m-order associative legendre polynomial; c_nmAnd S_nmIs the gravitational potential coefficient.

Further, the earth's internal mass distribution relationship is described by:

wherein s is a position vector of a point in the earth,and λ' are eachThe geocentric latitude and geocentric longitude corresponding to the point, rho(s) is the density of the point, d³s is the volume of the spot, δ_0mThe value of (a) depends on m, and the following relationship exists:

therefore, the gravity perturbation acceleration in the earth center fixed coordinate system can pass through V_nmAnd W_nmAnd calculating to obtain:

in the formula:

further, for high precision absolute kinetic models, consider the J of the dominant perturbation₂Item, then J₂The matrix form of the C-W equation of type is written as:

wherein the content of the first and second substances,

wherein, Δ r is a relative distance vector, s represents a sin function, c represents a cos function, r, u, i and Ω are respectively the sagittal diameter, orbital amplitude, orbital inclination and orbital elevation intersection declination of the target spacecraft, and μ is the earth gravity constant.

Further, after the relative navigation data and the absolute navigation data are calculated and output, the relative average orbit number is calculated according to the relative navigation data and the absolute navigation data.

In one possible implementation, calculating a relative average orbit number according to the relative navigation data and the absolute navigation data includes: calculating the position and the speed of the tracked spacecraft according to absolute navigation data, calculating the number of the close orbits of the tracked spacecraft according to the position and the speed, and calculating the average number of the orbits of the tracked spacecraft according to the number of the close orbits; and obtaining a relative average orbit root by adopting Kalman filtering UKF estimation according to the relative navigation data, the close orbit root and the average orbit root.

Specifically, in the solution provided in the embodiment of the present application, the relative average number of orbits is the difference between the average number of orbits of the tracked spacecraft and the average number of orbits of the target spacecraft, and is recorded as the differenceIn general, the relative orbit number is derived from absolute navigation and relative navigation, and the relative navigation has higher precision compared with the absolute navigation. A relative orbit element in a nonsingular form is estimated by adopting Kalman filtering (Unscented Kalman Filter, UKF), a relative position delta r and a relative speed delta v are used as random variables, and an absolute position r and an absolute speed v are used as determination variables. The specific process of calculating the relative average track number is shown in fig. 3.

In the whole task execution process, the tracking spacecraft is always in a configuration maintaining or rendezvous stage, and various orbit control instructions are triggered according to the relative navigation data output by the relative dynamic model. In order to ensure the rapid convergence of the UKF filter, two methods are adopted:

the method comprises the steps of 1, determining the sum of the pulse speed after the rail control and the speed before the rail control, calculating the state quantity after the rail control according to the sum of the speeds, updating the preset initial relative navigation data according to the state quantity, and outputting the updated relative navigation data.

And 2, controlling the kinematics model not to output relative navigation data within a preset time long range after orbit control until the UKF filter is ensured to achieve a preset convergence effect, and outputting the relative navigation data.

Specifically, in the scheme provided in the embodiment of the present application, the preset time period may be set according to an actual requirement, for example, the preset time period is 8000s, and is not limited herein. The method 1 updates the navigation data (relative navigation data and absolute navigation data) after orbit control, so that the filter has an initial value state quantity with small error in the initial stage of convergence, can quickly converge the filter, and belongs to an active control method; the method 2 ensures that the navigation data output is a convergence value by delaying the output of the kinematics model, and belongs to a passive control method. Specifically, the flow of the relative kinetic model output data is shown in fig. 4. In FIG. 4, an engine flag bit and a fire pulse command are received; judging whether the engine zone bit is 1 according to the engine pulse instruction, wherein the condition that the engine zone bit is 1 means that the engine executes rail control and can be automatically cleared after pulse action is executed; if not, outputting relative navigation data; if the number of the engine pulse vectors is 1, calculating and updating the state quantity after the orbit control according to the preset engine pulse vector, the position before the orbit control and the speed by using the relative dynamic model; acquiring timing data of the timer, judging whether the timing data is longer than a preset time (such as 8000s), and if not, acquiring the timing data of the timer again until the timing data is longer than the preset time; and if the time length is longer than the preset time length, outputting relative navigation data.

And 202, performing track recursion within a preset time period according to the relative navigation data, the absolute navigation data and the relative average track number to obtain track recursion data, and judging whether a preset rendezvous condition is met according to the track recursion data.

Specifically, in the scheme provided by the embodiment of the application, the UKF filter is used for relative navigation, so that linearization and calculation of Jacobian matrices are avoided, and the problems of precision and real-time performance of relative navigation are solved. In the autonomous orbit control strategy, orbit recursion is carried out for a plurality of times according to output navigation data (absolute navigation data and relative navigation data) of a dynamic model, the orbit recursion data is compared with a task configuration, and if a configuration threshold value is reached or intersection is required, an orbit control instruction is sent.

And 203, if so, generating an orbit control instruction set, calculating two control pulses of the tracking spacecraft at the initial time and the end time, and adjusting relative orbit parameters of the tracking spacecraft and the target spacecraft according to the orbit control instruction set and the two control pulses so as to enable the tracking spacecraft and the target spacecraft to carry out autonomous rendezvous.

In one possible implementation, two control pulses of the tracked spacecraft at the initial and end times are calculated, including: determining the difference of the initial semi-major axis between the tracking spacecraft and the target spacecraft, and calculating the difference of the expected semi-major axis according to the preset expected semi-major axis drift amount of each period; calculating to obtain a semi-major axis value which needs to be adjusted for tracking the spacecraft according to the difference between the initial semi-major axis and the expected semi-major axis; calculating to obtain a pulse value required to be adjusted by the tracking spacecraft according to the semi-long axis value and a preset perturbation equation, and determining a relative motion state transfer matrix according to the semi-long axis and the pulse value; calculating according to the relative motion state transition matrix, a preset relative motion model and positions of a preset initial moment and a preset terminal moment to obtain an initial speed and a terminal speed; and calculating to obtain two control pulses of the tracked spacecraft at the initial moment and the end moment according to the preset initial speed and the preset end speed as well as the initial speed and the end speed.

Specifically, in the solution provided in the embodiments of the present application, for autonomous rendezvous remote guidance, the target is approached by using the drift formed by the difference between the two-star semi-major axes. The relationship between the difference of the semimajor axes of the two stars and the drift amount of each period in the x-axis direction of the relative movement is as follows:

Δx＝-3πΔa (14)

wherein Δ a represents the difference between the semi-major axis of the target spacecraft and the semi-major axis of the tracking spacecraft. Suppose the drift amount expected by the drift approaching process per cycle is Deltax_expThe difference Δ a between the desired semi-major axes_expComprises the following steps:

Δa_exp＝Δx_exp/3π (15)

if the difference between the semi-major axes of the initial two stars is Δ a₀Then, the amount of the semi-major axis that needs to be adjusted for tracking the spacecraft is:

δa＝Δa₀-Δa_exp (16)

from the classical perturbation equation:

wherein a is a semi-major axis of the spacecraft orbit, e is the eccentricity of the spacecraft orbit, theta is the true proximal angle of the spacecraft orbit, r is the vector diameter of the spacecraft orbit, v is the speed of the spacecraft orbit, ft and fn are respectively the tangential force and the normal force borne by the spacecraft, and mu is the earth gravitational constant. Giving pulses Δ v tracking the spacecraft tuning semimajor axis δ a_x：

For autonomous rendezvous proximity guidance, a relative motion model is considered. The relative motion state transition equation can be obtained as follows:

wherein, X (t)₀) And V (t)₀) At a time t₀Relative position and velocity, X (t) and V (t) being relative position at time tAnd setting and speed, phi is a relative motion state transition matrix. Converting the above formula to the following form:

wherein G is a conversion form of the relative motion state transition matrix. If the initial time t₀Position X (t)₀) And terminal time t_fPosition X (t)_f) Having been determined, the required initial velocity V (t) can then be found₀) And terminal velocity V (t)_f). Again according to a given initial speed V' (t)₀) And terminal velocity V' (t)_f) Two control pulses at the initial time and the final time can be obtained.

Initial time control pulse:

ΔV₀＝V(t₀)-V′(t₀) (22)

end point timing control pulse:

ΔV_f＝V(t_f)-V′(t_f) (23)

further, in a possible implementation manner, the tracking command set includes: a track control instruction 1, a track control instruction 2, a track control instruction 3, a track control instruction 4, a track control instruction 5 and a track control instruction 6; the track control instruction 1 is used for controlling and changing the difference value of the relative half shafts to be a positive value; the orbit control instruction 2 is used for controlling and changing the relative half shaft difference value to be a negative value; the orbit control instruction 3 and the orbit control instruction 4 are respectively used for controlling and changing the orbit control relative eccentricity vectors at the initial time and the terminal time; the track control command 5 is used for controlling and changing the right ascension difference value of the relative ascending intersection point; and the track control command 6 is used for controlling and changing the relative inclination difference.

Further, in a possible implementation manner, the track control command 1, the track control command 2, the track control command 3, the track control command 4, the track control command 5, and the track control command 6 each include: instruction word, execution time and boot time.

Further, in one possible implementation, the tracking command set includes:

the tracking command set is represented by:

CIS＝[1，t₁，Δt₁；2，t₂，Δt₂；3，t₃，Δt₃；4，t₄，Δt₄；5，t₅，Δt₅；6，t₆，Δt₆；k，t_k，Δt_k]

wherein, 1, 2, 3, 4, 5 and 6 respectively represent instruction words of the track control instruction 1 to the track control instruction 6; t is t₁、t₂、t₃、t₄、t₅、t₆Respectively showing the execution time of the track control instruction 1 to the track control instruction 6; Δ t₁、Δt₂、Δt₃、Δt₄、Δt₅、Δt₆Respectively representing the startup duration corresponding to the track control instruction 1-the track control instruction 6; k represents an instruction word corresponding to the minimum execution time in the track control instructions 1 to 6; t is t_kRepresenting the minimum execution time in the track control commands 1 to 6; Δ t_kAnd the starting time corresponding to the minimum execution time instruction in the track control instructions 1 to 6 is shown.

Specifically, in the solution provided in the embodiments of the present application, the orbit control generally includes inter-satellite relative phase (Δ u), relative eccentricity vectors (Δ ex and Δ ey), relative tilt angle, and elevation crossing right ascension (Δ i and Δ Ω), and the relative phase is controlled by relative semi-major axis Δ a; considering the fact that the eccentricity vector of the tangential control is twice of the efficiency of the radial control, the fact that the orbit control engine is mostly installed on the + x surface or the-x surface, and the like, the eccentricity vector adopts the double-pulse tangential control. The control command set includes 6 tracking commands, namely a tracking command 1(CH1), a tracking command 2(CH2), a tracking command 3(CH3, a tracking command 4(CH4), a tracking command 5(CH5), and a tracking command 6(CH6), and the functions of the specific tracking commands are as follows:

1. CH1 changes the relative semimajor axis difference to be a positive value;

2. CH2 changes the relative semimajor axis difference to be a negative value;

3. the double pulses CH3 and CH4 modify the relative eccentricity vector, and CH3 is executed earlier than CH 4;

4. CH5 changes the relative elevation crossing right ascension difference;

5. CH6 changing the relative inclination angle difference;

further, each instruction consists of three parts: instruction word (i ═ 1, 2, …, 6), and execution time t_iAnd the starting length delta t of the engine_i. The Control Instruction Set (CIS) is defined as follows:

CIS＝[CH1；CH2；CH3；CH4；CH5；CH6；CHk]＝[1，t₁，Δt₁；2，t₂，Δt₂；3，t₃，Δt₃；4，t₄，Δt₄；5，t₅，Δt₅；6，t₆，Δt₆；k，t_k，Δt_k]

among these, CHk is an instruction having the smallest execution time among CH1 to CH 6.

Further, in a possible implementation manner, adjusting the relative orbit parameters of the tracking spacecraft and the target spacecraft according to the orbit control instruction set and the two control pulses includes: calculating a preset time delay extrapolation track parameter according to an input current track measurement parameter, and calculating to obtain track recursion data corresponding to a moment of calling an orbit control instruction set according to the extrapolation track parameter, wherein the preset time delay refers to the time delay from receiving navigation data to calling the orbit control instruction set; and calling the orbit control instruction set according to the orbit recursion data, adjusting relative orbit parameters of the tracking spacecraft and the target spacecraft according to the orbit control instruction set and two control pulses until any orbit control instruction reaches a preset boundary threshold value, and updating the orbit control instruction set.

In the scheme provided by the embodiment of the application, according to the duration T corresponding to thermal control, attitude control, ground control and preset time delay, the designed track recursion algorithm can input sigma according to the current orbit measurement₀Calculating extrapolated orbit sigma at time T_T(ii) a The extrapolation process does not need to accumulate intermediate data, and the extrapolation needs to adopt analytic calculation to call track recursion data corresponding to the track control instruction set moment. For the sake of understanding, the following description briefly describes the self-rendezvous process:

firstly, relative navigation or orbit determination software improves the input parameters of a controller according to an external sensor; relative control software extrapolates data at the time T according to the current external data and calls subprograms corresponding to CH 1-CH 6 in sequence, if a certain instruction reaches a boundary threshold value, an instruction set is updated, otherwise, the instruction set is maintained unchanged; transmitting the instruction of the minimum execution time in the instruction set to attitude control software, and judging whether to execute operations such as attitude maneuver and the like in advance; meanwhile, the command is transmitted to the rail-controlled engine, and after the command is executed, the engine needs to reversely send the command for executing the control currently and the shutdown identifier to the control software.

Furthermore, the temporary command (GCI) sent by the ground is not constrained by the extrapolated T moment, and is implemented immediately when the execution moment contained in the command is reached; the instruction is defined as GCI ═ k₁，t₁，k₂，t₂，Δt]Wherein k is₁Numbering the satellites to be controlled in the entire formation network, t₁For the moment of instruction upswing, k₂For the command to be controlled, t₂At is the instruction execution time, Δ t is the instruction execution length.

Furthermore, the meeting time sequence can be annotated on the ground and modified on line, and the new annotation time sequence takes effect at the moment T after the annotation; the meeting time sequence is defined asWherein, C_nTo configuration identifiers, t_nOpening time for the configuration; the control software embeds various configuration parameters and can read according to the identifier.

Further, in the process of tracing spacecraft rendezvous control, according to the requirement of task planning, the functions of configuration capture, configuration maintenance, rendezvous, fault module evacuation and the like are completed in different time periods. Based on the function design unified architecture software, 6 instructions are designed to sequentially complete different functions according to corresponding task identifiers.

Although the functions implemented by the 6 control instructions are similar, the respective priorities are different: CH3 and CH4 are twin instructions and the execution interval is half a orbital cycle; CH1 and CH2 are mutually exclusive instructions, i.e., they cannot occur simultaneously and the execution interval is at least greater than the preparation time T; CH5 and CH6 are also mutually exclusive instructions; in addition, the generation processes of CH1(CH2), CH3(CH4) and CH5(CH6) are independent from each other, so the following criteria need to be satisfied in the programming:

(1) after the instruction from CH1 to CH6 is generated, modification is not allowed until the instruction is executed, although the control boundary is gradually close to the threshold value in the period;

(2) the CH3 and CH4 commands are generated simultaneously, namely, the CH4 command needs to be generated according to the CH3 trigger condition;

(3) the phase change is caused during the adjustment of the eccentricity vector, namely CH3 and CH4 commands are forbidden to trigger CH1 or CH2 commands from generation to execution;

(4) the CH1(CH2) instruction allows the CH3(CH4) or CH5(CH6) instruction to be triggered from generation to execution;

(5) the CH5(CH6) allows the CH3(CH4) or CH1(CH2) instructions to be triggered from generation to execution.

Criterion (1) can be implemented by adding instruction trigger conditions: if | Δ t_kI < ε (where ε is a small amount set to avoid calculation errors, e.g. 1 × 10)^-5) Then the command-based tracking service process is called to determine whether the extrapolated orbit reaches the threshold, and the execution time t after the threshold is reached_kAnd execution amount Δ t_k(ii) a If | Δ t_kIf | > ε, the instruction has generated an instruction but has not yet executed, so the tracking process is directly ignored. Criterion (2) can be implemented by triggering the tracking command 3: t is t₄＝t₃+T_orb[ Delta ] t and [ 2 ]₄＝Δt₃(wherein T is_orbA track period). Criterion (3) can be implemented by instruction execution post-processing: feeding back an execution identifier by the engine after the execution of the CH3 is finished, and then clearing CH1, CH2 and CH3 in the instruction set; the execution of CH4 is completed by engine feedback execution identifier, followed by clearing of CH1, CH2, and CH4 within the instruction set. Criteria (4) and (5) may be implemented by instruction execution post-processing: after the engine feedback execution identifier is acquired, the CH1, CH2, CH5, and CH6 are emptied, respectively.

The method for calculating the execution time and the execution amount generated by each instruction track control service process in the configuration maintenance stage comprises the following steps:

CH1：wherein v is the current speed of the target satellite, a is the semi-major axis of the orbit of the target satellite, Da₁Desired value for the relative semi-major axis, Da₂Is the actual value relative to the semi-major axis, Da₁Given by an empirical value of 15 m; considering that the formation maneuver is executed in China as much as possible, the execution position is selected to be executed at a position with a dimensional argument equal to 30 degrees, the time required by the satellite to reach the position is calculated according to the current dimensional argument and the orbit extrapolation time T is added, and the execution time recorded as CH1 is T₁；

CH 2: the calculation method is the same as CH 1;

CH3：wherein, | De | ═ Δ e_t- Δ e |; the execution position isWhereinCalculating the argument u of the arrival latitude according to the argument of the current dimension₁The required time plus the extrapolated time of the orbit T, denoted as CH3, has an execution time T₃；

CH4：The execution position is u₂＝u₁+π，u₁For the latitude argument calculated for CH3, the arrival u is calculated from the current dimension argument₂The required time plus the extrapolated time of the orbit T, denoted as CH4, has an execution time T₄；

CH5：Δv₅Obtaining a target star current orbit inclination angle, wherein the target star current orbit inclination angle is v × D Ω × sin i; the execution position is latitude argumentCalculating the time required for reaching the position and adding the upper rail according to the amplitude angle of the current dimensionExtrapolation time T, denoted as CH5, with execution time T₅；

CH6：Δv₆Vxxi Di |; the execution position is latitude argument u is 0, the time needed for reaching the position is calculated according to the current dimension argument and the track extrapolation time T is added, and the execution time recorded as CH6 is T₆。

The method for calculating the execution time and the execution amount generated by each instruction track control service process in the rendezvous stage comprises the following steps:

CH1：wherein v is the current speed of the target satellite, a is the semi-major axis of the orbit of the target satellite, Da₁Desired value for the relative semi-major axis, Da₂Is the actual value relative to the semi-major axis, Da₁Given as an empirical value of 25 m; considering that the formation maneuver is executed in China as much as possible, the execution position is selected to be executed at a position with a dimensional argument equal to 30 degrees, the time required by the satellite to reach the position is calculated according to the current dimensional argument and the orbit extrapolation time T is added, and the execution time recorded as CH1 is T₁；

CH 2: the calculation method is the same as CH 1;

CH3：wherein, | De | ═ Δ e_t- Δ e |; the execution position isWhereinCalculating the argument u of the arrival latitude according to the argument of the current dimension₁The required time plus the extrapolated time of the orbit T, denoted as CH3, has an execution time T₃；

CH4：The execution position is u₂＝u₁+π，u₁Latitude argument calculated for CH3, based onThe current dimension argument calculates the arrival u₂The required time plus the extrapolated time of the orbit T, denoted as CH4, has an execution time T₄；

CH5：Δv₅V × D Ω × sin i, where i is the current orbital inclination of the target star; the execution position is latitude argumentCalculating the time required for reaching the position and the track extrapolation time T according to the amplitude angle of the current dimension, and recording the execution time as the execution time T of CH5₅；

CH6：Δv₆Vxxi Di |; the execution position is latitude argument u is 0, the time needed for reaching the position is calculated according to the current dimension argument and the track extrapolation time T is added, and the execution time recorded as CH6 is T₆。CH4：The execution position is u₂＝u₁+π，u₁For the latitude argument calculated for CH3, the arrival u is calculated from the current dimension argument₂The required time plus the extrapolated time of the orbit T, denoted as CH4, has an execution time T₄；

CH5：Δv₅Obtaining a target star current orbit inclination angle, wherein the target star current orbit inclination angle is v × D Ω × sin i; the execution position is latitude argumentCalculating the time required for reaching the position and the track extrapolation time T according to the amplitude angle of the current dimension, and recording the execution time as the execution time T of CH5₅；

CH6：Δv₆Vxxi Di |; the execution position is latitude argument u is 0, the time needed for reaching the position is calculated according to the current dimension argument and the track extrapolation time T is added, and the execution time recorded as CH6 is T₆。

Further, in the scheme provided in the embodiment of the present application, if it is determined according to the track recurrence data that the preset rendezvous condition is not met, the relative navigation data is calculated and output again according to the relative dynamics model and the absolute navigation data is calculated according to the absolute dynamics model until the track recurrence data determines that the preset rendezvous condition is met.

In order to facilitate understanding of the beneficial effects of the method for controlling the in-orbit autonomous intersection of the micro/nano satellite provided in the embodiment of the present application, the following description is made in an exemplary manner.

For example, consider that the tracking star and the target star are both in the same orbital plane, 90 degrees out of phase, and the long-range transfer period is 10 orbital periods. The tracking star initial orbit elements are as follows: semi-major axis a of the track_s6735000m, eccentricity e_sTrack inclination i ═ 0_s0.8727, right ascension omega_s5.9341, argument of perigee ω_s0, latitude argument M_s1.6484. The initial orbit elements of the target star are as follows: semi-major axis a of the track_t6735000m, eccentricity e_tTrack inclination i ═ 0_t0.8727, right ascension omega_t5.9341, argument of perigee ω_t0, latitude argument M_t＝0.0784。

By applying the micro-nano satellite in-orbit autonomous intersection control simulation analysis method provided by the embodiment of the application, the following results are obtained: fig. 5 gives a schematic diagram of the spacing between the tracking spacecraft star and the target spacecraft. After 10 minutes when the remote first maneuvering time is in the input state, the maneuvering amount is 62.508 meters per second, the second maneuvering time is 48 minutes and 4 seconds on the same year, the same month, the next day and the fourth day, and the maneuvering amount is 39.072 meters per second. The two maneuvers consumed 1.0154 kilograms of fuel in total.

In the scheme provided by the embodiment of the application, relative navigation data are calculated and output according to the relative dynamics model, absolute navigation data are calculated and obtained according to the absolute dynamics model, and relative average orbit number is calculated according to the relative navigation data and the absolute navigation data; in other words, in the scheme provided by the embodiment of the application, absolute and relative dynamics models are respectively adopted, so that the spacecraft orbit change intersection butt joint control precision under the actual shooting condition is improved.

Referring to fig. 6, an embodiment of the present application provides a computer device, including:

a memory 601 for storing instructions for execution by at least one processor;

a processor 602 for executing instructions stored in a memory to perform the method as described in fig. 2.

As will be appreciated by one skilled in the art, embodiments of the present application may be provided as a method, system, or computer program product. Accordingly, the present application may take the form of an entirely hardware embodiment, an entirely software embodiment or an embodiment combining software and hardware aspects. Furthermore, the present application may take the form of a computer program product embodied on one or more computer-usable storage media (including, but not limited to, disk storage, optical storage, and the like) having computer-usable program code embodied therein.

The present application is described with reference to flowchart illustrations and/or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the application. It will be understood that each flow and/or block of the flow diagrams and/or block diagrams, and combinations of flows and/or blocks in the flow diagrams and/or block diagrams, can be implemented by computer program instructions. These computer program instructions may be provided to a processor of a general purpose computer, special purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, create means for implementing the functions specified in the flowchart flow or flows and/or block diagram block or blocks.

These computer program instructions may also be stored in a computer-readable memory that can direct a computer or other programmable data processing apparatus to function in a particular manner, such that the instructions stored in the computer-readable memory produce an article of manufacture including instruction means which implement the function specified in the flowchart flow or flows and/or block diagram block or blocks.

These computer program instructions may also be loaded onto a computer or other programmable data processing apparatus to cause a series of operational steps to be performed on the computer or other programmable apparatus to produce a computer implemented process such that the instructions which execute on the computer or other programmable apparatus provide steps for implementing the functions specified in the flowchart flow or flows and/or block diagram block or blocks.

It will be apparent to those skilled in the art that various changes and modifications may be made in the present application without departing from the spirit and scope of the application. Thus, if such modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is intended to include such modifications and variations as well.