Spacecraft formation configuration decoupling control method
By employing the E/I vector method for spacecraft formation control and utilizing radial, track, and normal relative motion equations, various control strategies were designed. This approach solved the problems of error accumulation and high fuel consumption in long-duration, wide-range flight using the CW equation, achieving efficient and stable formation control.
Patent Information
- Application Number
- CN202511961083.5
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Filing Date
- 2025-12-24
- Publication Date
- 2026-03-20
- Estimated Expiration
- 2045-12-24
AI Technical Summary
Existing spacecraft formation control methods based on the CW equation suffer from severe accumulation of approximate errors, high fuel consumption, and difficulty in decoupling control laws during long-duration, wide-range flights, making it difficult to meet the requirements of high-precision formation missions.
The E/I vector method is used for spacecraft formation control. By calculating the relative orbital elements, the radial, track, and normal relative motion equations are obtained. Combined with J2 perturbation, single-pulse normal, double-pulse track, and multi-cycle track control strategies are designed to achieve stable control of the relative orbital elements.
Effective separation of orbit shape and orbital plane control reduces control frequency, propellant consumption, and improves long-term operational stability and economy, achieving singularity-free and well-decoupled formation control.
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Figure CN121386848B_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present application relates to a spacecraft formation configuration decoupling control method, and belongs to the technical field of satellite formation control. BACKGROUND
[0002] Among the existing spacecraft formation control methods, the most mainstream one is the spacecraft formation control method based on the CW equation (Clothoid-Wirth). The CW equation is an equation describing relative motion derived under the assumption of near-circular and the condition that the distance between two satellites is much smaller than the distance from the center of the Earth. The spacecraft formation control method based on the CW equation is suitable for small-scale and short-term tasks, but in long-term and large-scale formation flight, there is a serious accumulation of approximation errors, which needs to be frequently corrected and has the defect of high fuel consumption.
[0003] In contrast, the E / I vector method (eccentricity / inclination vector method) avoids the above problems while maintaining the essential characteristics of orbital dynamics. The E / I vector method is intuitive in geometric sense and can effectively separate the shape of the orbit from the control of the orbital plane, thereby simplifying the control law design for formation keeping and reconstruction. In addition, since the E / I vector evolves relatively smoothly over time, the control frequency can be reduced, reducing propellant consumption and improving the stability and economy of long-term operation. Therefore, the E / I vector method has unique advantages of no singularity, good decoupling, high fuel efficiency and suitability for long-term tasks in formation flight control. SUMMARY
[0004] In view of the problem of serious accumulation of approximation errors and high fuel consumption in the spacecraft formation control method based on the CW equation, the present application provides a spacecraft formation configuration decoupling control method.
[0005] The spacecraft formation configuration decoupling control method of the present application comprises:
[0006] The relative orbital elements of the satellite and the primary star are calculated based on the orbital elements of the satellite and the primary star;
[0007] Based on the relative orbital elements, the radial, trace and normal relative motion equations of the satellite and the primary star are obtained; and considering the J2 perturbation, the modified radial, trace and normal relative motion equations are obtained;
[0008] The velocity pulse is determined by the dynamic equation between the relative orbital elements and the velocity pulse;
[0009] The required change amount of the relative orbit root number is determined according to the relative orbit root number expected value and the relative orbit root number current value, the required change amount of the corresponding relative orbit inclination vector, the required change amount of the relative semi-major axis, the required change amount of the relative eccentricity vector and the required change amount of the relative angular distance are determined according to the required change amount of the relative orbit root number, the single-pulse method is selected to control the relative orbit inclination vector according to the dynamic equation between the relative orbit root number and the velocity pulse, the double-pulse trace direction pulse strategy is selected to control the relative semi-major axis and the relative eccentricity vector, the multi-period trace direction control strategy is selected to control the relative angular distance according to the corrected radial, trace direction and normal relative motion equation, and finally the relative orbit root number reaches the relative orbit root number expected value.
[0010] The method of the application is based on E / I vector (eccentricity vector and inclination vector) for spacecraft formation control, the eccentricity vector describes the orbit shape and the perigee direction, and the inclination vector describes the direction of the orbit plane. As the slowly varying variables of the orbit, the two vectors change smoothly over time, and there is no singularity problem when the eccentricity or inclination approaches zero, thus providing a more stable parameterization for formation control.
[0011] Compared with the traditional method, the E / I vector method makes up for many shortcomings. First, the control method based on the relative motion model (such as CW or YA equation) is suitable for small-scale and short-term tasks, but the approximation error accumulates seriously in long-term and large-scale formation flight, which needs frequent correction, resulting in high fuel consumption. Second, the method of directly using classical orbital elements is easy to fail when e or i approaches zero, and the coupling between orbital elements is strong, making it difficult to decouple and implement the control law. Third, the method based on the average perturbation theory is suitable for long-term analysis, but it often ignores short-period effects, and the calculation is complex and difficult to meet the needs of high-precision formation tasks. In contrast, the E / I vector method maintains the essential characteristics of orbital dynamics while avoiding the above problems. It is intuitive in geometric sense and can effectively separate the control of orbit shape and orbit plane, thus simplifying the control law design of formation keeping and reconstruction. In addition, since the E / I vector evolves relatively smoothly over time, the control frequency can be reduced, reducing propellant consumption and improving the stability and economy of long-term operation. Therefore, the E / I vector method has unique advantages of no singularity, good decoupling, high fuel efficiency and suitability for long-term tasks in formation flight control. BRIEF DESCRIPTION OF DRAWINGS
[0012] Figure 1 is a control strategy flowchart of the spacecraft formation decoupling control method described in the application;
[0013] Figure 2 is a schematic diagram of the relative motion of the primary star and the secondary star in the radial and trace direction planes;
[0014] Figure 3 is a schematic diagram of relative motion of the primary star and the satellite in the radial and normal planes;
[0015] Figure 4 is a flight control flowchart used in actual engineering;
[0016] Figure 5 is a RTN direction trajectory diagram in the simulation experiment;
[0017] Figure 6 is a RT direction trajectory diagram in the simulation experiment;
[0018] Figure 7 is a trajectory diagram of the deviation root number ;
[0019] Figure 8 is a trajectory diagram of the deviation root number ;
[0020] Figure 9 is a trajectory diagram of the deviation root number ;
[0021] Figure 10 is a trajectory diagram of the deviation root number ;
[0022] Figure 11 is a control strategy schematic diagram in the task process;
[0023] Figure 12 is a required velocity increment schematic diagram in the task process. DETAILED DESCRIPTION
[0024] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative work fall within the scope of protection of the present application.
[0025] Specifically, the present application provides a spacecraft formation configuration decoupling control method, comprising: Figure 1
[0026] The relative orbit root number of the satellite and the primary star is calculated according to the orbit root number of the primary star and the satellite;
[0027] The radial, trace and normal relative motion equations of the satellite and the primary star are obtained based on the relative orbit root number, and the corrected radial, trace and normal relative motion equations are obtained by considering the J2 perturbation;
[0028] The velocity pulse is determined by a dynamic equation between the relative orbital elements and the velocity pulse;
[0029] The required change amount of the relative orbital elements is determined according to the relative orbital element expected value and the relative orbital element current value, and the required change amount of the corresponding relative orbit inclination vector, the required change amount of the relative semi-major axis, the required change amount of the relative eccentricity vector and the required change amount of the relative angular distance are determined according to the required change amount of the relative orbital elements; the single pulse method normal control strategy is selected to control the relative orbit inclination vector in combination with the dynamic equation between the relative orbital elements and the velocity pulse; the double pulse trace direction pulse strategy is selected to control the relative semi-major axis and the relative eccentricity vector; the multi-period trace direction control strategy is selected to control the relative angular distance in combination with the corrected radial, trace direction and normal relative motion equation, so that the relative orbital elements finally reach the relative orbital element expected value.
[0030] The present embodiment is based on the E / I vector for spacecraft formation control, and has clear and simple logic, low calculation requirement and easy engineering implementation. After the maneuvering transfer, before approaching and capturing the target, the platform and the target to be captured will perform a period of formation flight to further observe the target and perform self-checking. Therefore, the relative state of the two stars needs to be accurately controlled to achieve stable flying. The spacecraft formation control based on the E / I vector decouples the formation control into out-of-plane motion and in-plane motion for control. The present embodiment uses the near-circular non-equatorial Gauss variation equation to design a flying control strategy using only normal and tangential pulses to perform out-of-plane and in-plane formation control. In modeling, the influence of the earth's spheroid and atmospheric perturbation is considered and compensated in control, so that the algorithm model is relatively accurate. After design, the algorithm structure is relatively simple and can effectively control different states.
[0031] The specific process is as follows: when the ground obtains the target orbital elements, the platform obtains the relative orbital elements of the two stars, judges whether control is needed and what kind of control is needed. After generating the control instruction, the calculation pulse is applied to the specified position to achieve stable formation flying. In the stable formation flying stage, the single control pulse can be less than 0.1 m / s, and the control interval can reach 1 day, which can meet the formation requirements of the two stars before capture.
[0032] Further, in the geocentric inertial coordinate system, the orbital elements of the main star are defined as:
[0033] (1),
[0034] In the formula, a is the semi-major axis of the main star, is the orbital elements of the main star, is the semi-major axis of the main star, is the argument of latitude of the main star, is the argument of latitude of the main star, the mean anomaly of the primary star, the X-axis component of the eccentricity vector of the primary star, the Y-axis component of the eccentricity vector of the primary star, the eccentricity of the primary star, the inclination of the orbit of the primary star, the right ascension of the ascending node of the primary star;
[0035] the relative orbital elements of the secondary star with respect to the primary star are:
[0036] (2)
[0037] wherein is the relative orbital elements, is the relative semi-major axis, is the relative argument of latitude, is the X-axis component of the relative eccentricity vector, is the Y-axis component of the relative eccentricity vector, is the X-axis component of the relative inclination vector, is the Y-axis component of the relative inclination vector, is the semi-major axis of the secondary star, is the argument of latitude of the secondary star, is the right ascension of the ascending node of the secondary star, is the X-axis component of the eccentricity vector of the secondary star, is the Y-axis component of the eccentricity vector of the secondary star, is the inclination of the orbit of the secondary star.
[0038] Neglecting the effects of perturbations and performing first-order linearization, the radial, transverse and normal relative motion equations of the secondary star with respect to the primary star are obtained in the LVLH coordinate system of the primary star based on the relative orbital elements:
[0039] (3),
[0040] wherein is the radial position perturbation, is the transverse position perturbation, is the normal position perturbation.
[0041] In formula (3), the right side of the equation is the relative state quantity at time 0, and the left side is the state quantity corresponding to time u. is a variable, and the rest are relative state quantities at time 0, and the left side is the state quantity corresponding to time u. Figure 2 and Figure 3 describe such relative motion. Figure 2 in which denotes the radial direction, denotes the transverse direction; Figure 3 in which denotes the normal direction.
[0042] In the stable formation flying process, the target star is the primary star and the tracking star is the secondary star. The primary star does not perform orbit maneuvering, and only the secondary star is maneuvered to control the relative state. In the control process, a virtual primary star can be selected, and the state between the virtual primary star and the real primary star meets the task requirements, mainly the interstellar distance requirement. The advantage of this is that the control target is to control the secondary star to the position of the virtual primary star, even if the relative deviation is less than a certain threshold.
[0043] The radial, trace, and normal relative motion equations after J2 perturbation correction are considered as follows:
[0044] (4),
[0045] where is the relative eccentricity vector, is the phase angle of the relative eccentricity vector , is the derivative of , is the initial azimuth of the primary star, is the perturbation coefficient.
[0046] (5),
[0047] (6),
[0048] where is the perturbation coefficient, is the radius of the earth, is an intermediate variable, , and can be taken as 1 under the near-circular condition.
[0049] .
[0050] The following gives the dynamic equation describing the relationship between the velocity pulse and the relative state, which guides the design of the velocity pulse, i.e., the required control quantity.
[0051] The dynamic equation between the relative orbit root and the velocity pulse is as follows:
[0052] (7),
[0053] where is the trace velocity pulse increment applied by the secondary star, is the orbit angular velocity amplitude, is the radial velocity pulse, is the application position of the velocity pulse applied by the secondary star, is the normal velocity pulse.
[0054] Equation (7) is usually called Gauss Variation Equation (GVE) of near-circular non-equatorial orbit. The left side of the equation is the variation of relative orbit elements, and the control objective is to make the relative orbit elements deviation less than a threshold value close to zero due to the strategy of virtual primary. Therefore, the variation of relative orbit elements should be the negative value of the actual deviation, so as to realize positive and negative offset and control the deviation of relative orbit elements within the threshold range; the right side is the velocity impulse applied at In practical tasks, in general, the ignition time is less than or equal to the order of 100 seconds, which can be considered as a pulse velocity increment.
[0055] Analyzing equation (7), except , the rest of the state variables can be changed instantaneously by the velocity pulse, The change of needs to be accumulated in time. In addition, in the relative orbit elements, only the relative orbit inclination vector is related to the normal velocity pulse, and it is known from the relative motion equation that the normal motion is only related to the relative orbit inclination vector , so the normal motion can be decoupled and controlled, which is called out-of-plane motion. The radial and trace motions are called in-plane motions, which need to be coupled and controlled by using radial and trace velocity pulses. The specific control strategy is given below.
[0056] First, calculate the required variation of relative orbit elements, and then calculate the required pulse according to the relative dynamics equation. The required variation of relative orbit elements is denoted as
[0057] (8),
[0058] where is the expected value of relative orbit elements, is the current value of relative orbit elements.
[0059] Judge the required variation of relative orbit elements :
[0060] If , no control; otherwise, judge the relative orbit inclination vector :
[0061] If , use the single-pulse normal control strategy to control the relative orbit inclination vector ; otherwise, judge the relative semi-major axis and the relative eccentricity vector :
[0062] If or , use the double-pulse trace pulse strategy to control the relative semi-major axis and the relative eccentricity vector control; otherwise, control the relative angular distance control is performed:
[0063] If , the multi-period trace control strategy is used to control the relative angular distance , otherwise the control is ended;
[0064] wherein is the control deviation threshold value, is the required change amount of the relative orbit inclination vector, is the relative orbit inclination vector deviation threshold value, is the required change amount of the relative semi-major axis, is the relative semi-major axis deviation threshold value, is the required change amount of the relative eccentricity vector, is the relative eccentricity vector deviation threshold value, is the required change amount of the relative angular distance, is the relative angular distance deviation threshold value, .
[0065] The method for controlling the relative inclination vector using the single-pulse normal control strategy is:
[0066] (9),
[0067] (10),
[0068] (11),
[0069] wherein is the required change amount of the relative orbit inclination vector X-axis component, is the required change amount of the relative orbit inclination vector Y-axis component.
[0070] Due to the coupling effect of radial and tangential pulses, in-plane motion control is more complex. The following method is given for controlling in-plane motion only using tangential pulses.
[0071] The method for controlling the relative semi-major axis and the relative eccentricity vector using the double-pulse trace pulse strategy is:
[0072] If or , and :
[0073] (12),
[0074] (13),
[0075] (14),
[0076] wherein is the amplitude of the trace impulse one, is an intermediate variable, is an intermediate variable, is the amplitude of the trace impulse two, is an intermediate variable, is the phase angle of the relative eccentricity vector , is the position of the trace impulse one, is the position of the trace impulse two; wherein the position of the trace impulse one may be optional;
[0077] wherein:
[0078] (15),
[0079] , , (16),
[0080] (17);
[0081] wherein is the X-axis component of the required change of the relative eccentricity vector, is the Y-axis component of the required change of the relative eccentricity vector,
[0082] if or and :
[0083] (18),
[0084] (19),
[0085] (20),
[0086] (21).
[0087] When the relative orbit inclination vector , the relative semi-major axis and the relative eccentricity vector are all less than the corresponding set threshold, i.e. the normal and radial positions both satisfy the control index, the relative angular distance is controlled.
[0088] The method for controlling the relative angular distance by using the multi-period trace control strategy is:
[0089] (22),
[0090] (twenty three),
[0091] In the formula To adjust the amplitude of the first track velocity pulse during track separation, To adjust the amount of change required relative to the semi-major axis during track separation, To adjust the expected value relative to the semi-major axis during trace separation, The current relative semi-major axis of the primary star and the secondary star. The amount of change in the mean latitude argument required to separate the paths of the primary and secondary stars. The orbital period of the main star;
[0092] The principle is to adjust the radial distance by utilizing the difference in orbital angular velocity caused by the difference in the semi-major axis. In fact, due to the relative orbital inclination vector... and relative eccentricity vector It has been controlled within the threshold, as can be seen from the relative kinematic equations. Mainly composed of Therefore, this strategy can be used for adjustment. When considering J2 perturbation and atmospheric drag:
[0093] (twenty four),
[0094] (25)
[0095] In the formula The primary star's mean-dimensional argument used in guidance calculations. The sub-star's planar argument used for guidance calculations. This represents the current X-axis component of the relative orbital tilt vector. This is the relative ballistic coefficient. For the atmospheric density of the orbital environment, Main star velocity; Main star ballistic coefficient, This is the atmospheric drag coefficient. To ensure effective windward area, For satellite quality;
[0096] The last two terms in Equation (24) are the additional drift caused by J2 and atmospheric perturbation during the drift over time. Equation (24) takes into account the additional drift and compensates for it during control.
[0097] In the formula For sub-star ballistic coefficients;
[0098] When the ballistic coefficients of the two satellites are inconsistent, the effect of atmospheric drag during long-term operation cannot be ignored.
[0099] After running for several cycles, the trajectory separation is adjusted to the required range, at which time the trajectory pulse is applied, adjusting , avoiding the divergence of the trajectory separation.
[0100] (26),
[0101] In the formula is the amplitude of the second trajectory velocity pulse when adjusting the trajectory separation.
[0102] The natural trajectory drift time is two cycles, so nT is 4 , that is, after running two tracks, a tangential pulse is applied at the same position, adjusting , ending the trajectory drift.
[0103] The specific implementation process is shown in Figure 1 Each time the inter-satellite communication is obtained, the relative orbit root number is obtained, and according to the specific situation of the relative root number, the corresponding strategy is selected for control.
[0104] Figure 4 A flight control flowchart used in an actual engineering is given. After the program is started, some variables are initialized / kept first, and then it is judged whether there is an orbit control command according to the flag quantity, if there is, the command is executed, otherwise the guidance instruction calculation is performed.
[0105] For the guidance calculation branch, the orbit root number of the spacecraft in the formation is first obtained, which can be obtained from the inter-satellite communication or from the ground. After obtaining the orbit root number, it can be judged that the current formation configuration is compared with the set threshold to judge whether to adjust and maintain. If each relative root number meets the requirement, that is, no control is required, other tasks can be executed, otherwise the guidance instruction is calculated, and after the calculation is completed and stable, the flag quantity is assigned to the required orbit control and the control instruction exists.
[0106] For the guidance instruction execution branch, the state of the orbit control thruster is first judged, and according to the state of the thruster, the spacecraft can be respectively entered into the star service takeover, thruster work and orbit control preparation branch. For the orbit control preparation branch, due to the algorithm design, no guidance instruction needs to be generated immediately, and the attitude adjustment time is reserved, so first it is judged whether to enter the ignition interval (that is, to reach the ignition position), after reaching the ignition position, the attitude adjustment flag is enabled, and the attitude is adjusted to meet the required thrust direction; after reaching the ignition position, the thruster flag is enabled, and the thruster is modulated to work until completion.
[0107] The application scenario of the method of the application is:
[0108] 1. Spacecraft formation flying: It is suitable for spacecraft formation flying control, which can autonomously determine and calculate the guidance strategy on the satellite, and realize autonomous spacecraft maneuvering by adjusting the attitude and thruster switch state to meet the formation flying task requirements.
[0109] 2. Autonomous rendezvous and docking of spacecraft: It is suitable for autonomous rendezvous and docking tasks between spacecraft. Before spacecraft rendezvous and docking, parking state checking is required, etc. The formation flying control method of the present application can realize parking along the flight path direction, fully meeting the task requirements.
[0110] 3. On-orbit service and maintenance: It is suitable for on-orbit service and maintenance. In this task, close proximity to the target is required, so remote parking is also required for related inspection before subsequent tasks can be carried out. Therefore, the formation flying control of the present application is also suitable for this task.
[0111] Simulation experiment:
[0112] The simulation scene setting is shown in Table 1, wherein sat1 is the main star, which does not perform orbit maneuvering, and sat2 is the sub-star, which performs orbit maneuvering to make the configuration meet the set requirements. The two stars are in low earth orbit, and there are differences in semi-major axis, eccentricity, orbit inclination and dimension pitch angle, so various control strategies and effects can be verified. For simulation, the specific control target is to completely control the sub-star to the position of the main star, so that the positions of the two stars coincide. It is worth noting that when there is no main star, this control can be regarded as orbit adjustment or orbit maintenance of a single star. The control result is shown in Figures 1 to 8 .
[0113] Table 1
[0114]
[0115] In Table 1 is the satellite true anomaly.
[0116] Figure 5 and Figure 6 shows the relative trajectory of the sub-star in the RTN coordinate system (LVLH coordinate system) of the main star during the control process. Figures 7 to 10 shows the change of the deviation root number during the task process. Figure 5 In the RTN coordinate system, R represents the radial direction, pointing from the earth center to the spacecraft; N represents the normal direction, pointing to the positive normal direction of the spacecraft orbit plane; T represents the trace direction, constituting a right-handed system with R and N. Figure 7 In the RTN coordinate system, R represents the radial direction, pointing from the earth center to the spacecraft; N represents the normal direction, pointing to the positive normal direction of the spacecraft orbit plane; T represents the trace direction, constituting a right-handed system with R and N.
[0117] Figure 11 shows the control strategy during the task process, strategy 1 represents single-pulse normal control, corresponding to multi-period trace control strategy; strategy 2 and strategy 3 represent in-plane control, corresponding to double-pulse trace pulse strategy, wherein strategy 2 corresponds to Case 3 corresponds to the strategy 3 Case 4 represents the adjustment of the separation of the trajectory, corresponding to the multi-period trajectory control strategy; and the strategy 0 represents no control. Figure 12 The speed increment required in the task process is shown. Figure 12 The total speed increment is shown.
[0118] It can be seen that after about 5e4 seconds, the configuration initialization has been basically completed, that is, the E / I vector is controlled to be within the threshold range of 50m. In the initialization process, the in-plane motion adjustment and the separation of the trajectory adjustment are mainly performed, and the total speed pulse required is about 1.2m / s. In the subsequent task time, the orbit maintenance is mainly performed, that is, the maintenance is kept at the main star position, and the adjustment of the separation of the trajectory is mainly performed. In the period of 5e4 to 2.5e5, only two adjustments of the separation of the trajectory are performed, and the rest of the time does not require control, and the speed increment required is far less than 0.1m / s.
[0119] As can be seen from the above, in the process of the accompanying flight task, the satellite formation flight can be realized by using the method of the application, and the precision can meet the set threshold. The task process can be roughly divided into two stages of configuration initialization and configuration maintenance, and the fuel consumption is mainly concentrated in the process of configuration initialization, and the fuel required for configuration maintenance is extremely small.
[0120] In summary, the method of the application has the following advantages:
[0121] 1. Spacecraft formation control based on E / I vector: the kinematics and dynamics of spacecraft formation flight are modeled by E / I vector, the motion is decomposed into in-plane motion and out-of-plane motion, and the out-of-plane motion control strategy, the in-plane accompanying ellipse control strategy and the in-plane trajectory separation control strategy are designed according to the motion characteristics.
[0122] 2. Control logic and flow design: based on the design of the in-plane and out-of-plane control strategy, the control logic is proposed, that is, the out-of-plane motion is controlled first, then the in-plane accompanying ellipse motion is controlled, and finally the in-plane trajectory separation motion is controlled. Based on the logic, the specific control flow is designed, which can realize relatively efficient and stable spacecraft formation control.
[0123] 3. Multi-period trajectory separation adjustment strategy: a multi-period trajectory separation adjustment strategy is designed by using the natural drift of the angular velocity difference, and by appropriately relaxing the adjustment time, the required velocity pulse can be greatly reduced to complete the low-energy trajectory separation adjustment. In addition, by modifying the expected companion flying position and setting the corresponding threshold, various companion flying conditions can be realized, including but not limited to only out-of-plane motion, in-plane fixed-point keeping; only in-plane elliptical flying; only in-plane fixed-point keeping, etc.
[0124] 4. Actual engineering case process: combined with the design of a specific task of a certain type, the full autonomous companion flying task on the satellite can be realized. According to the characteristics of the type, a process for realizing the companion flying task can be designed to realize autonomous inter-satellite communication, control demand judgment, guidance instruction generation and execution and other operations.
[0125] The method of the present application provides a decoupled direct method: by modeling the E / I vector, the motion is decoupled into in-plane / out-of-plane motion, which reduces the control complexity, and by further decomposing the in-plane motion, i.e. companion flying elliptical control and trajectory separation control, the control complexity is further reduced, and the engineering practicability is improved.
[0126] Decouple control of each motion in multiple tracks: compared with completing multiple controls in one track or in a shorter time, the method of the present application can simply and efficiently complete formation control according to the logic of first controlling out-of-plane motion, then controlling in-plane elliptical companion flying, and finally controlling trajectory separation. The coupling effect caused by multiple pulse control in a single track, the inability to achieve the expected control target, and the difficulty of the control algorithm which is not conducive to engineering implementation are avoided.
[0127] The method of the present application has strong engineering applicability: the algorithm calculation is simple, can normally run on a computer board with a calculation frequency of 1Hz, and through reasonable division of code levels, the algorithm logic is rigorous, and in the case of lower computing power, the corresponding control effect can also be realized, thereby reducing the system hardware cost and structural complexity, so that the algorithm has strong engineering applicability.
[0128] In the method of the present application, the main star refers to the satellite that does not perform orbit maneuver in the formation task, and the sub-star performs orbit maneuver around the main star, so that the double-star configuration meets the set requirements.
[0129] The sub-star refers to the satellite that performs orbit maneuver in the formation task, and performs orbit maneuver around the main star, so that the double-star configuration meets the set requirements.
[0130] While the invention has been described herein with reference to specific embodiments, it should be understood that these embodiments are merely examples of the principles and applications of the invention. Therefore, it should be understood that many modifications can be made to the exemplary embodiments, and other arrangements can be designed without departing from the spirit and scope of the invention as defined by the appended claims. It should be understood that different dependent claims and features described herein can be combined in ways different from those described in the original claims. It is also understood that features described in conjunction with individual embodiments can be used in other described embodiments.
Claims
1. A spacecraft formation configuration decoupling control method, characterized in that, include: The relative orbital elements between the primary star and the primary star are calculated based on the orbital elements of the primary star and the secondary star. Based on the relative orbital elements, the radial, track, and normal relative motion equations between the secondary star and the primary star are obtained; Considering the J2 perturbation, we obtain the corrected radial, trace, and normal relative motion equations; The velocity pulse is determined by the dynamic equation relating the relative orbital elements and the velocity pulse. Based on the expected value and current value of the relative orbital elements, the required change in the relative orbital elements is determined. From the required change in the relative orbital elements, the required changes in the relative orbital inclination vector, the relative semi-major axis, the relative eccentricity vector, and the relative angular distance are determined. Combining the dynamic equations between the relative orbital elements and velocity pulses, a single-pulse normal control strategy is selected to control the relative orbital inclination vector. A dual-pulse track pulse strategy is selected to control the relative semi-major axis and the relative eccentricity vector. Combining the corrected radial, track, and normal relative motion equations, a multi-cycle track control strategy is selected to control the relative angular distance, ultimately enabling the relative orbital elements to reach the expected value. The change required for the relative orbital elements is expressed as... : （8）， In the formula This represents the expected value of the relative orbital elements. This is the current value of the relative orbital elements; The amount of change required for relative orbital elements Make a judgment: like No control; otherwise, the relative orbital tilt vector Make a judgment: like A single-pulse normal control strategy is used to control the relative orbital inclination vector. Control is required; otherwise, the relative semi-major axis is affected. and relative eccentricity vector Make a judgment: like or A dual-pulse tracing pulse strategy is used to target the relative semi-major axis. and relative eccentricity vector Control it; otherwise, adjust the relative angular distance. Make a judgment: like A multi-cycle trajectory control strategy is adopted to control the relative angular distance. Perform control; otherwise, terminate control. In the formula To control the deviation threshold, This is the required change in the relative orbital inclination vector. The relative orbital inclination vector deviation threshold. This is the amount of change required relative to the semi-major axis. The relative semi-major axis deviation threshold. The required change in the relative eccentricity vector. Relative eccentricity vector deviation threshold This is the required change in relative angular distance. The relative angular distance deviation threshold. ; The method for controlling the relative semi-major axis and relative eccentricity vector using a dual-pulse trajectory pulse strategy is as follows: like or ,and : （12）， （13）， （14）， In the formula The amplitude of the first trace pulse. As an intermediate variable, As an intermediate variable, The amplitude of the second trace pulse. As an intermediate variable, The relative eccentricity vector phase angle, The position of the first trace pulse. The position of the second trace pulse; The magnitude of the orbital angular velocity. The semi-major axis of the main star; in: （15）， ， ， （16）， （17）； In the formula The required change in the relative eccentricity vector is the X-axis component. The required change in the relative eccentricity vector is the Y-axis component. The X-axis component of the relative eccentricity vector. The relative eccentricity vector is the Y-axis component; like or ,and : （18）， （19）， （20）， （21）。 2. The spacecraft formation configuration decoupling control method according to claim 1, characterized in that, The orbital elements of a primary star are defined as: （1）， In the formula The orbital elements of the main star, The mean latitude and argument of the main star, Angular distance from the perigee of the primary star The main star's near-point angle, The X-axis component of the principal star's eccentricity vector. The Y-axis component of the principal star's eccentricity vector. Eccentricity of the principal star The orbital inclination of the primary star. Right ascension of the ascending node of the principal star; The relative orbital elements between the secondary star and the primary star are: （2） In the formula The relative orbital elements, For the relative semi-major axis, The relative angular distance The X-axis component of the relative orbital inclination vector. The relative orbital inclination vector Y-axis component, For the semi-major axis of the sub-star, The argument of the secondary star's mean latitude. The right ascension of the ascending node of the child star, The x-axis component of the eccentricity vector of the child star. The Y-axis component of the eccentricity vector of the child star. The inclination of the sub-star's orbit.
3. The spacecraft formation configuration decoupling control method according to claim 2, characterized in that, Based on the aforementioned relative orbital elements, the radial, path, and normal relative motion equations between the secondary star and the primary star are obtained: （3）， In the formula This is the radial position perturbation. For the trajectory position perturbation, This is the normal position perturbation.
4. The spacecraft formation configuration decoupling control method according to claim 3, characterized in that, After considering the J2 perturbation correction, the radial, path, and normal relative motion equations are: （4）， In the formula The relative eccentricity vector, The relative eccentricity vector phase angle, for The derivative, The initial mean latitude and argument of the main star, The perturbation coefficient; （5）， （6）， In the formula The perturbation coefficient is... For the Earth's radius, As an intermediate variable, .
5. The spacecraft formation configuration decoupling control method according to claim 4, characterized in that, The dynamic equation between relative orbital elements and velocity pulses is: （7）， In the formula The incremental trajectory velocity pulse applied to the sub-star It is a radial velocity pulse. The location at which the velocity pulse is applied to the sub-star. It is a normal velocity pulse.
6. The spacecraft formation configuration decoupling control method according to claim 5, characterized in that, The method for controlling the relative tilt vector using a single-pulse normal control strategy is as follows: （9）， （10）， （11）， In the formula The X-axis component represents the required change in the relative orbital inclination vector. The Y-axis component represents the required change in the relative orbital inclination vector.
7. The spacecraft formation configuration decoupling control method according to claim 6, characterized in that, The method for controlling the relative angular distance using a multi-cycle trajectory control strategy is as follows: （22）， （23）， In the formula To adjust the amplitude of the first track velocity pulse during track separation, To adjust the amount of change required relative to the semi-major axis during track separation, To adjust the expected value relative to the semi-major axis during trace separation, The current relative semi-major axis of the primary star and the secondary star. The amount of change in the mean latitude argument required to separate the paths of the primary and secondary stars. The orbital period of the main star; （24）， （25）， In the formula The primary star's mean-dimensional argument used in guidance calculations. The sub-star's planar argument used for guidance calculations. This represents the current X-axis component of the relative orbital tilt vector. This is the relative ballistic coefficient. For the atmospheric density of the orbital environment, Main star velocity; Main star ballistic coefficient, This is the atmospheric drag coefficient. To ensure effective windward area, For satellite quality; （26）， In the formula To adjust the amplitude of the second trajectory velocity pulse during trajectory separation.