Remote guidance method and device based on anchor point integrated common rail and direct rendezvous

Abstract

The present invention relates to the field of rendezvous guidance technology, and particularly to a long-range guidance method and device for integrating co-orbital and direct rendezvous based on an anchor point. Among them, the method is applied to the guidance system of a tracking spacecraft, and includes: obtaining the position parameters of a target anchor point, a co-orbital guidance terminal, and a rendezvous and docking terminal; obtaining a flag quantity used to characterize whether the tracking spacecraft is faulty, so as to determine a long-range guidance terminal from the co-orbital guidance terminal and the rendezvous and docking terminal; when the tracking spacecraft reaches the target anchor point, performing impulse guidance based on the position parameters of the target anchor point and the long-range guidance terminal, so that the tracking spacecraft reaches the long-range guidance terminal and completes the long-range guidance of the tracking spacecraft. In this solution, by reasonably designing the target anchor point, it is determined whether the tracking spacecraft is faulty before the tracking spacecraft reaches the target anchor point, and the long-range guidance terminal is determined according to whether there is a fault, so that the tracking spacecraft can take into account the autonomous long-range guidance in both normal and faulty situations.

Description

Technical Field

[0001] The embodiments of the present invention relate to the technical field of rendezvous guidance, and particularly to a long-range guidance method and device for integrating co-orbital and direct rendezvous based on anchor points. Background Art

[0002] With the continuous development of aerospace technology, various tracking spacecraft with different functions are continuously launched into space orbits. After entering the space orbit, through the long-range guidance of the tracking spacecraft, the tracking spacecraft is made to fly in the orbital direction of the target spacecraft to complete its own mission by flying in the same orbit as the target spacecraft or flying in formation.

[0003] However, various failures may occur after the tracking spacecraft enters the space orbit, such as the failure of the solar panel deployment. Then, it is necessary to first perform a rendezvous and docking with the target spacecraft to repair the failure of the tracking spacecraft. However, the existing spacecraft guidance methods cannot take into account these two situations to achieve the autonomous long-range guidance mission of the tracking spacecraft.

[0004] Therefore, there is an urgent need for a long-range guidance method for integrating co-orbital and direct rendezvous based on anchor points. Summary of the Invention

[0005] In order to solve the problem that the existing spacecraft guidance methods cannot take into account the long-range guidance in both normal and failure situations, the embodiments of the present invention provide a long-range guidance method and device for integrating co-orbital and direct rendezvous based on anchor points.

[0006] In a first aspect, the embodiments of the present invention provide a long-range guidance method for integrating co-orbital and direct rendezvous based on anchor points, which is applied to the guidance system of a tracking spacecraft. The method includes:

[0007] Obtaining the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal;

[0008] Obtaining a flag quantity used to represent whether the tracking spacecraft is faulty, so as to determine the long-range guidance terminal from the co-orbital guidance terminal and the rendezvous and docking terminal;

[0009] When the tracking spacecraft reaches the target anchor point, performing pulse guidance based on the position parameters of the target anchor point and the position parameters of the long-range guidance terminal, so that the tracking spacecraft reaches the long-range guidance terminal and completes the long-range guidance of the tracking spacecraft.

[0010] In a second aspect, the embodiments of the present invention further provide a long-range guidance device for integrating co-orbital and direct rendezvous based on anchor points, which is applied to the guidance system of a tracking spacecraft and includes:

[0011] An obtaining unit, configured to obtain the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal;

[0012] A determination unit, configured to obtain a flag indicating whether the tracking spacecraft is faulty, so as to determine a remote guidance terminal from the co-orbital guidance terminal and the rendezvous and docking terminal;

[0013] A guidance unit, configured to perform pulse guidance based on the position parameters of the target anchor point and the position parameters of the remote guidance terminal when the tracking spacecraft reaches the target anchor point, so that the tracking spacecraft reaches the remote guidance terminal and completes the remote guidance of the tracking spacecraft.

[0014] An embodiment of the present invention provides a remote guidance method and device for integrating co-orbital and direct rendezvous based on an anchor point, which is applied to the guidance system of a tracking spacecraft. First, obtain the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal; then, obtain a flag indicating whether the tracking spacecraft is faulty, so as to determine a remote guidance terminal from the co-orbital guidance terminal and the rendezvous and docking terminal; finally, when the tracking spacecraft reaches the target anchor point, perform pulse guidance based on the position parameters of the target anchor point and the position parameters of the remote guidance terminal, so that the tracking spacecraft reaches the remote guidance terminal and completes the remote guidance of the tracking spacecraft. In this solution, by inserting a target anchor point, it is determined whether the tracking spacecraft is faulty before the tracking spacecraft reaches the target anchor point, and the remote guidance terminal is determined according to whether there is a fault, so that the tracking spacecraft can take into account autonomous remote guidance in both normal and faulty situations. Description of the Drawings

[0015] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the description of the embodiments or the prior art. Obviously, the following drawings are some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0016] Figure 1 is a flowchart of a remote guidance method for integrating co-orbital and direct rendezvous based on an anchor point provided by an embodiment of the present invention;

[0017] Figure 2 is a schematic diagram of a spatial position in a rendezvous and docking coordinate system provided by an embodiment of the present invention;

[0018] Figure 3 is a hardware architecture diagram of a computing device provided by an embodiment of the present invention;

[0019] Figure 4 is a structural diagram of a remote guidance device for integrating co-orbital and direct rendezvous based on an anchor point provided by an embodiment of the present invention. Detailed Embodiments

[0020] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are some, but not all, of the embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts fall within the scope of protection of the present invention.

[0021] As described above, after the tracking spacecraft separates from the rocket and enters the space orbit, if everything is normal with the tracking spacecraft after entering the space orbit, during the remote guidance phase, the tracking spacecraft needs to be guided to the co-orbital guidance terminal and then transferred to co-orbital flight; if the tracking spacecraft malfunctions after entering the space orbit, such as a failure in the deployment of the solar panels, etc., and it cannot be solved through remote control, then during the remote guidance phase, the tracking spacecraft needs to be guided to the rendezvous and docking terminal and then transferred to the short-range autonomous rendezvous phase to complete the rendezvous and docking of the tracking spacecraft and the target spacecraft, so as to use the on-orbit maintenance function of the target spacecraft to repair the faults of the tracking spacecraft.

[0022] However, the existing remote guidance methods for spacecraft cannot take into account both of these situations to achieve the autonomous remote guidance mission of the tracking spacecraft.

[0023] To solve the above technical problems, the inventor can consider designing a target anchor point during the remote guidance phase, determining whether the tracking spacecraft malfunctions before reaching the target anchor point, and determining which of the co-orbital guidance terminal and the rendezvous and docking terminal is the remote guidance terminal according to whether there is a malfunction. Then, the tracking spacecraft can perform pulse guidance based on the position parameters of the target anchor point and the position parameters of the remote guidance terminal to complete the remote guidance mission. It can be seen that this solution can take into account the autonomous remote guidance missions for two different destinations of rendezvous and docking and co-orbital flight.

[0024] The following describes the specific implementation manners of the above concept.

[0025] Please refer to Figure 1 , the embodiments of the present invention provide a remote guidance method for integrating co-orbital and direct rendezvous based on an anchor point, which is applied to the guidance system of a tracking spacecraft. The method includes:

[0026] Step 100, obtaining the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal;

[0027] Step 102, obtaining a flag quantity used to characterize whether the tracking spacecraft malfunctions, so as to determine the remote guidance terminal from the co-orbital guidance terminal and the rendezvous and docking terminal;

[0028] Step 104: When the tracking spacecraft reaches the target anchor point, perform pulse guidance based on the position parameters of the target anchor point and the position parameters of the remote guidance terminal, so that the tracking spacecraft reaches the remote guidance terminal, and complete the remote guidance of the tracking spacecraft.

[0029] In the embodiment of the present invention, first, obtain the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal; then, obtain a flag indicating whether the tracking spacecraft is faulty, so as to determine the remote guidance terminal from the co-orbital guidance terminal and the rendezvous and docking terminal; finally, when the tracking spacecraft reaches the target anchor point, perform pulse guidance based on the position parameters of the target anchor point and the position parameters of the remote guidance terminal, so that the tracking spacecraft reaches the remote guidance terminal, and complete the remote guidance of the tracking spacecraft. In this solution, by inserting the target anchor point, it is determined whether the tracking spacecraft is faulty before the tracking spacecraft reaches the target anchor point, and the remote guidance terminal is determined according to whether there is a fault, so that the tracking spacecraft can take into account the autonomous remote guidance in both normal and faulty situations.

[0030] Regarding step 100:

[0031] In some embodiments, the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal are determined in the following manner:

[0032] S1. Determine the spatial position of the target spacecraft that co-orbits and rendezvouses with the tracking spacecraft;

[0033] S2. Based on the spatial position of the target spacecraft, determine the spatial position model of the target anchor point;

[0034] S3. Based on the spatial position of the target spacecraft and the spatial position model of the target anchor point, determine the spatial position models of the co-orbital guidance terminal and the rendezvous and docking terminal;

[0035] S4. Based on the spatial position models of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal, determine the initial values of the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal;

[0036] S5. Based on the initial values of the position parameters, respectively determine the execution times of the two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in the two cases where the co-orbital guidance terminal is used as the remote guidance terminal and the rendezvous and docking terminal is used as the remote guidance terminal;

[0037] S6. Respectively determine whether the execution times of the corresponding two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in each case meet the energy constraint and trajectory safety requirements;

[0038] S7. If it is not satisfied in at least one case, based on the spatial position models of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal, adjust the initial values of the position parameters, and jump to execute S5 until the two guidance pulses corresponding to the two cases both meet the energy constraint and trajectory safety requirements, then obtain the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal.

[0039] In this embodiment, according to the spatial position of the target spacecraft, the spatial positions of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal are modeled in sequence. By setting the initial values of the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal, respectively determine the execution times of the two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in the two cases where the co-orbital guidance terminal is used as the far guidance terminal and the rendezvous and docking terminal is used as the far guidance terminal, and respectively determine whether the execution times of the two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in each case meet the energy constraint and trajectory safety requirements. If at least one case does not meet the requirements, adjust the initial values of the position parameters, and use the adjusted initial values of the position parameters to re-determine the execution times of the two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in each case until the guidance pulses in the two cases both meet the energy constraint and trajectory safety requirements, then obtain the final position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal. It can be seen that in this embodiment, under the condition of fully considering multiple constraint conditions such as energy constraint and flight trajectory safety requirements, the positions of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal are reasonably designed, improving the reliability of the tracking spacecraft's orbit insertion.

[0040] In some embodiments, the spatial position model of the target anchor point in step S2 is:

[0041]

[0042] Wherein, h p and h a Satisfy the following formula:

[0043]

[0044] In the formula, θ0 is the phase angle difference between the target anchor point and the target spacecraft, and the target anchor point is located behind the target spacecraft, θ min and θ max Are respectively the lower limit value and the upper limit value of the engineering constraint, x and z are the relative positions of the rendezvous and docking terminal on the x-axis and z-axis in the rendezvous and docking coordinate system, a T Is the orbital semi-major axis of the target spacecraft, h p Is the perigee altitude of the orbit where the target anchor point is located, h a Is the apogee altitude of the orbit where the target anchor point is located, Δap and Δa a are the perigee orbital altitude difference and apogee orbital altitude difference between the target anchor point and the rendezvous and docking terminal respectively. θ1 is the phase angle interval of the tracking spacecraft from the target anchor point to the position where the last guidance pulse is implemented before reaching the rendezvous and docking terminal, and θ2 is the phase angle interval of the tracking spacecraft from the position of the last guidance pulse to the rendezvous and docking terminal.

[0045] Reference Figure 2 Schematic diagram of the spatial position in the rendezvous and docking coordinate system. O is the center of mass of the target spacecraft. The Z-axis points to the center of mass of the earth. The Y-axis (not marked in the figure) is perpendicular to the OZ-axis (perpendicular to the paper surface) and points to the orbital angular velocity direction. The OX-axis, OZ-axis, and OY-axis form a right-handed coordinate system. P0 is the target anchor point, S0 is the co-orbital guidance terminal, and A0 is the rendezvous and docking terminal. After the tracking spacecraft enters the orbit, it reaches the target anchor point P0 through multiple guidances. If a fault of the tracking spacecraft is found before reaching the target anchor point P0, the rendezvous and docking terminal A0 is used as the far guidance terminal. After reaching the rendezvous and docking terminal A0, the tracking spacecraft is guided through short-range control to rendezvous and dock with the target spacecraft via B3, B2, and B1. If the tracking spacecraft is normal before reaching the target anchor point P0, the co-orbital guidance terminal S0 is used as the far guidance terminal. After reaching the co-orbital guidance terminal S0, it flies in the same orbit or flies in formation with the target spacecraft.

[0046] In this embodiment, θ1 generally can take half of the orbital period, that is, θ1 = 180°, and generally 0° ≤ θ2 ≤ 180°. θ min and θ max are engineering constraints from the feasibility of pulse implementation, and generally can take θ max = θ inorb θ inorb is the phase angle difference between the tracking spacecraft and the target spacecraft when the tracking spacecraft enters the orbit. Moreover, the satisfaction conditions of h p and h a make the distance between the target anchor point and the orbit entry point of the tracking spacecraft relatively far, which can ensure sufficient fault handling time. It can be seen that the design of the target anchor point fully considers the orbital control feasibility, TT&C condition constraints, and fault handling time constraints, and can improve the reliability of the tracking spacecraft entering the orbit.

[0047] In some embodiments, the spatial position model of the co-orbital guidance terminal in step S3 is:

[0048] The co-orbital guidance terminal is located behind the target spacecraft. The phase angle difference between the co-orbital guidance terminal and the target spacecraft is θ s , and the altitude difference between the co-orbital guidance terminal and the target spacecraft is Δa; where, θ s and Δa satisfy the following formula:

[0049]

[0050] In the formula, θ0 is the phase angle difference between the target anchor point and the target spacecraft, and θ s is the phase angle difference between the co-orbital guidance terminal and the target spacecraft, θ1 is the phase angle interval of the tracking spacecraft from the target anchor point to the position where the last guidance pulse is implemented before the rendezvous and docking terminal, Δa is the altitude difference between the co-orbital guidance terminal and the target spacecraft, and a T is the semi-major axis of the orbit of the target spacecraft, and a C is the semi-major axis of the orbit of the tracking spacecraft, and f(Δa) is the relationship function between the orbit altitude difference and the phase difference between the target spacecraft and the tracking spacecraft.

[0051] In this embodiment, f(Δa) is a relationship function set according to the space environment and mission requirements. Furthermore, through the conditional formula of θ s and Δa, the position of the co-orbital guidance terminal can be designed more reasonably, the reliability of orbit injection can be improved, and the engineering practicability is strong.

[0052] In some embodiments, the space position model of the rendezvous and docking terminal in step S3 is:

[0053]

[0054] In the formula, ω T is the orbital angular velocity of the target spacecraft, x, y, and z are the relative positions of the rendezvous and docking terminal on the x-axis, y-axis, and z-axis in the rendezvous and docking coordinate system, and v x , v x and v x are the relative velocities of the rendezvous and docking terminal in the x-axis direction, y-axis direction, and z-axis direction in the rendezvous and docking coordinate system, respectively.

[0055] In some embodiments, step S4 may include:

[0056] Based on the space position model of the target anchor point, determine the initial values of the position parameters of the target anchor point; wherein, the initial values of the position parameters of the target anchor point include θ0, x, θ1, θ2, Δa p and Δa a ;

[0057] Based on θ0, θ1, and the space position model of the co-orbital guidance terminal, determine the initial values of the position parameters of the co-orbital guidance terminal; wherein, the initial values of the position parameters of the co-orbital guidance terminal include θ s and Δa;

[0058] Based on x and the space position model of the rendezvous and docking terminal, determine the initial values of the position parameters of the rendezvous and docking terminal; wherein, the initial values of the position parameters of the rendezvous and docking terminal include z.

[0059] In this embodiment, the position parameters of the target anchor point refer to the parameters to be solved in the spatial position model of the target anchor point. These are related to the position of the target anchor point. The same applies to the position parameters of the common-rail guidance terminal and the rendezvous and docking terminal.

[0060] In some embodiments, step S5: Based on the initial values of the position parameters, respectively determine the execution times of the two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in the two cases of taking the common-rail guidance terminal as the remote guidance terminal and taking the rendezvous and docking terminal as the remote guidance terminal, which may include:

[0061] Taking the rendezvous and docking terminal as the remote guidance terminal, based on the initial values of the position parameters of the target anchor point, the initial values of the position parameters of the rendezvous and docking terminal, and the combined correction guidance algorithm, determine the latitude argument corresponding to the execution times of the two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in this first case;

[0062] Taking the common-rail guidance terminal as the remote guidance terminal, based on the initial values of the position parameters of the target anchor point, the initial values of the position parameters of the common-rail guidance terminal, and the combined correction guidance algorithm, determine the latitude argument corresponding to the execution times of the two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in this second case.

[0063] In this embodiment, the rendezvous and docking terminal A0 can be taken as the remote guidance terminal first. This is the first case. Then, using the combined correction guidance algorithm of the semi-major axis, eccentricity, and latitude argument, the latitude arguments u1, u2 corresponding to the execution times of the two guidance pulses, and the semi-major axis change amounts da1, da2 corresponding to the guidance pulses can be solved. Then, taking the common-rail guidance terminal S0 as the remote guidance terminal, this is the second case. Using the combined correction guidance algorithm of the semi-major axis, eccentricity, and latitude argument, the latitude arguments u1 * , u2 * , and the semi-major axis change amounts da1 * , da2 * .

[0064] After obtaining the guidance pulses in the two cases, according to step S6, first judge whether u1, u2, da1, and da2 in the first case meet the trajectory safety requirements and energy constraints (propellant consumption); then judge whether u1 * , u2 * , and da1 * , da2 *Whether the trajectory safety requirements and energy constraints (propellant consumption) are met. If not met in at least one case, based on the spatial position models of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal, adjust the initial values of the position parameters and jump to execute S5 until the two corresponding guidance pulses in both cases meet the energy constraints and trajectory safety requirements, and then obtain the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal.

[0065] In the embodiment of the present invention, the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal need to be calculated in advance. After the tracking spacecraft enters space, the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal that have been stored in the guidance system of the tracking spacecraft can be directly obtained to perform the processes of step 102 and step 104. Therefore, this solution can quickly generate the guidance schemes for two branches and reduce the calculation time of the tracking spacecraft in space.

[0066] Regarding step 102:

[0067] In the embodiment of the present invention, step 102 includes:

[0068] When the flag represents a fault of the tracking spacecraft, determine the rendezvous and docking terminal as the far guidance terminal;

[0069] When the flag represents the normal state of the tracking spacecraft, determine the co-orbital guidance terminal as the far guidance terminal.

[0070] In this embodiment, after the tracking spacecraft enters the orbit, it reaches the target anchor point P0 through multiple guidances. There is sufficient time between the tracking spacecraft entering the orbit and reaching the target anchor point P0. Once a fault occurs during this period, the fault can be repaired through remote control. When reaching the target anchor point P0, if the fault has been repaired or the tracking spacecraft has been normal since entering the orbit, the flag is the default value, then the co-orbital guidance terminal S0 will be determined as the far guidance terminal; when reaching the target anchor point P0 and the fault has not been repaired yet, the ground can inject and modify the flag to determine the rendezvous and docking terminal A0 as the far guidance terminal for subsequent guidance pulse calculation.

[0071] Regarding step 104:

[0072] When the common-rail guidance terminal S0 is determined as the remote guidance terminal in step 102, after the tracking spacecraft reaches the target anchor point, its guidance system will perform impulse guidance based on the position parameters of the target anchor point and the common-rail guidance terminal S0 to guide the tracking spacecraft to the common-rail guidance terminal S0, and then perform common-rail flight or accompanying flight with the target spacecraft; when the rendezvous and docking terminal A0 is determined as the remote guidance terminal in step 102, after the tracking spacecraft reaches the target anchor point, its guidance system will perform impulse guidance based on the position parameters of the target anchor point and the rendezvous and docking terminal A0 to guide the tracking spacecraft to the rendezvous and docking terminal A0, that is, the remote guidance is completed, and then the short-range control is entered to enable the tracking spacecraft to perform rendezvous and docking with the target spacecraft step by step.

[0073] As Figure 3 , Figure 4 shown, an embodiment of the present invention provides a remote guidance device for integrating common-rail and direct rendezvous based on an anchor point. The device embodiment can be implemented by software, or by hardware or a combination of software and hardware. In terms of the hardware level, as Figure 3 shown, it is a hardware architecture diagram of a computing device where a remote guidance device for integrating common-rail and direct rendezvous based on an anchor point provided by an embodiment of the present invention is located. In addition to Figure 3 the processor, memory, network interface, and non-volatile memory shown, the computing device where the device is located in the embodiment usually may further include other hardware, such as a forwarding chip responsible for processing packets, etc. Taking software implementation as an example, as Figure 4 shown, as a logically meaningful device, it is formed by the CPU of its computing device reading the corresponding computer program in the non-volatile memory into the memory and running. A remote guidance device for integrating common-rail and direct rendezvous based on an anchor point provided by this embodiment is applied to the guidance system of a tracking spacecraft and includes:

[0074] An acquisition unit 401, configured to acquire the position parameters of the target anchor point, the common-rail guidance terminal, and the rendezvous and docking terminal;

[0075] A determination unit 402, configured to acquire a flag quantity for characterizing whether the tracking spacecraft is faulty, so as to determine a remote guidance terminal from the common-rail guidance terminal and the rendezvous and docking terminal;

[0076] A guidance unit 403, configured to perform impulse guidance based on the position parameters of the target anchor point and the remote guidance terminal when the tracking spacecraft reaches the target anchor point, so that the tracking spacecraft reaches the remote guidance terminal and completes the remote guidance of the tracking spacecraft.

[0077] In an embodiment of the present invention, it further includes a calculation unit 404, configured to calculate the position parameters of the target anchor point, the common-rail guidance terminal, and the rendezvous and docking terminal in the acquisition unit 401;

[0078] The calculation unit 404 is configured to execute:

[0079] S1. Determine the spatial position of the target spacecraft that is in co - orbital flight and rendezvous - docking with the tracking spacecraft;

[0080] S2. Based on the spatial position of the target spacecraft, determine the spatial position model of the target anchor point;

[0081] S3. Based on the spatial position of the target spacecraft and the spatial position model of the target anchor point, determine the spatial position models of the co - orbital guidance terminal and the rendezvous - docking terminal;

[0082] S4. Based on the spatial position models of the target anchor point, the co - orbital guidance terminal, and the rendezvous - docking terminal, determine the initial values of the position parameters of the target anchor point, the co - orbital guidance terminal, and the rendezvous - docking terminal;

[0083] S5. Based on the initial values of the position parameters, respectively determine the execution times of the two guidance pulses and the semi - major axis change amount corresponding to each guidance pulse in two cases: when the co - orbital guidance terminal is used as the far - range guidance terminal and when the rendezvous - docking terminal is used as the far - range guidance terminal;

[0084] S6. Respectively determine whether the execution times of the corresponding two guidance pulses and the semi - major axis change amount corresponding to each guidance pulse in each case meet the energy constraint and trajectory safety requirements;

[0085] S7. If at least one case does not meet the requirements, based on the spatial position models of the target anchor point, the co - orbital guidance terminal, and the rendezvous - docking terminal, adjust the initial values of the position parameters, and jump to execute S5 until the two guidance pulses corresponding to the two cases both meet the energy constraint and trajectory safety requirements, then obtain the position parameters of the target anchor point, the co - orbital guidance terminal, and the rendezvous - docking terminal.

[0086] In an embodiment of the present invention, the spatial position model of the target anchor point in the calculation unit 404 is:

[0087]

[0088] Wherein, h p and h a Satisfy the following formula:

[0089]

[0090] In the formula, θ0 is the phase - angle difference between the target anchor point and the target spacecraft, and the target anchor point is located behind the target spacecraft, θ min and θ max Are respectively the lower limit value and the upper limit value of the engineering constraint, x and z are the relative positions of the rendezvous - docking terminal on the x - axis and the z - axis in the rendezvous - docking coordinate system, aT is the semi-major axis of the target spacecraft's orbit, h p is the perigee altitude of the orbit where the target anchor point is located, h a is the apogee altitude of the orbit where the target anchor point is located, Δa p and Δa a are respectively the perigee orbit altitude difference and the apogee orbit altitude difference between the target anchor point and the rendezvous and docking terminal. θ1 is the phase angle interval of the tracking spacecraft from the target anchor point to the last guidance pulse before the rendezvous and docking terminal, and θ2 is the phase angle interval of the tracking spacecraft from the last guidance pulse to the rendezvous and docking terminal.

[0091] In one embodiment of the present invention, in the calculation unit 404, the spatial position model of the co-orbital guidance terminal is:

[0092] The co-orbital guidance terminal is located behind the target spacecraft, and the phase angle difference between the co-orbital guidance terminal and the target spacecraft is θ s , and the altitude difference between the co-orbital guidance terminal and the target spacecraft is Δa; where θ s and Δa satisfy the following formula:

[0093]

[0094] In the formula, θ0 is the phase angle difference between the target anchor point and the target spacecraft, θ s is the phase angle difference between the co-orbital guidance terminal and the target spacecraft, θ1 is the phase angle interval of the tracking spacecraft from the target anchor point to the last guidance pulse before the rendezvous and docking terminal, Δa is the altitude difference between the co-orbital guidance terminal and the target spacecraft, a T is the semi-major axis of the target spacecraft's orbit, a C is the semi-major axis of the tracking spacecraft's orbit, and f(Δa) is the relationship function between the orbit altitude difference and the phase difference between the target spacecraft and the tracking spacecraft.

[0095] In one embodiment of the present invention, the spatial position model of the rendezvous and docking terminal in the calculation unit 404 is:

[0096]

[0097] In the formula, ω T is the orbital angular velocity of the target spacecraft, x, y, and z are respectively the relative positions of the rendezvous and docking terminal on the x-axis, y-axis, and z-axis in the rendezvous and docking coordinate system, v x , v x and v x are respectively the relative velocities of the rendezvous and docking terminal in the x-axis direction, y-axis direction, and z-axis direction in the rendezvous and docking coordinate system.

[0098] In one embodiment of the present invention, when the calculation unit 404 executes step S5, it is used for:

[0099] Taking the rendezvous and docking terminal as the remote guidance terminal, based on the initial value of the position parameter of the target anchor point, the initial value of the position parameter of the rendezvous and docking terminal, and the combined correction guidance algorithm, determine the latitude argument corresponding to the execution times of the two guidance pulses in this first case and the semi-major axis change amount corresponding to each guidance pulse.

[0100] Taking the co-orbital guidance terminal as the remote guidance terminal, based on the initial value of the position parameter of the target anchor point, the initial value of the position parameter of the co-orbital guidance terminal, and the combined correction guidance algorithm, determine the latitude argument corresponding to the execution times of the two guidance pulses in this second case and the semi-major axis change amount corresponding to each guidance pulse.

[0101] In an embodiment of the present invention, the determining unit 402 is configured to execute:

[0102] When the flag represents a failure of the tracking spacecraft, determine the rendezvous and docking terminal as the remote guidance terminal;

[0103] When the flag represents the normal state of the tracking spacecraft, determine the co-orbital guidance terminal as the remote guidance terminal.

[0104] It can be understood that the structure schematically shown in the embodiments of the present invention does not constitute a specific limitation on a remote guidance device for the integration of co-orbital and direct rendezvous based on an anchor point. In other embodiments of the present invention, a remote guidance device for the integration of co-orbital and direct rendezvous based on an anchor point may include more or fewer component units than shown in the figure, or combine certain component units, or split certain component units, or have different arrangements of component units. The component units shown in the figure can be implemented in hardware, software, or a combination of software and hardware.

[0105] Regarding the information interaction, execution process, etc. among the various modules within the above-mentioned device, since they are based on the same concept as the method embodiments of the present invention, the specific content can be referred to the description in the method embodiments of the present invention, and will not be elaborated here.

[0106] An embodiment of the present invention further provides a computing device, including a memory and a processor. A computer program is stored in the memory. When the processor executes the computer program, it implements a remote guidance method for the integration of co-orbital and direct rendezvous based on an anchor point in any embodiment of the present invention.

[0107] An embodiment of the present invention further provides a computer-readable storage medium. A computer program is stored on the computer-readable storage medium. When the computer program is executed by a processor, it causes the processor to execute a remote guidance method for the integration of co-orbital and direct rendezvous based on an anchor point in any embodiment of the present invention.

[0108] Specifically, a system or device equipped with a storage medium can be provided. On this storage medium, software program codes for implementing the functions of any one of the above embodiments are stored, and the computer (or CPU or MPU) of the system or device is made to read and execute the program codes stored in the storage medium.

[0109] In this case, the program codes read from the storage medium itself can implement the functions of any one of the above embodiments. Therefore, the program codes and the storage medium storing the program codes constitute a part of the present invention.

[0110] Examples of the storage medium for providing program codes include floppy disks, hard disks, magneto-optical disks, optical disks (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), magnetic tapes, non-volatile memory cards, and ROMs. Optionally, the program codes can be downloaded from a server computer via a communication network.

[0111] In addition, it should be clear that not only can the functions of any one of the above embodiments be realized by executing the program codes read by the computer, but also by making the operating system or the like operating on the computer based on the instructions of the program codes to complete part or all of the actual operations.

[0112] In addition, it can be understood that the program codes read from the storage medium are written into the memory provided in the expansion board inserted into the computer or into the memory provided in the expansion module connected to the computer. Subsequently, based on the instructions of the program codes, the CPU or the like installed on the expansion board or the expansion module is made to execute part and all of the actual operations, thereby realizing the functions of any one of the above embodiments.

[0113] It should be noted that in this article, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the term "comprising", "including" or any other variant thereof is intended to cover non-exclusive inclusion, such that a process, method, article or device comprising a series of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such process, method, article or device.

[0114] Those of ordinary skill in the art can understand that all or part of the steps of implementing the above method embodiments can be completed by hardware related to program instructions. The aforementioned program can be stored in a computer-readable storage medium. When the program is executed, it performs the steps including those of the above method embodiments. The aforementioned storage medium includes various media that can store program codes, such as ROM, RAM, magnetic disks, or optical discs.

[0115] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit them. Although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that they can still modify the technical solutions described in the foregoing embodiments or perform equivalent replacements for some of the technical features. These modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the various embodiments of the present invention.

Claims

1. A remote guidance method based on anchor - point integrated common - rail and direct rendezvous, characterized in that, A guidance system applied to a tracking spacecraft, the method comprising: Obtaining the position parameters of a target anchor point, a co-orbital guidance terminal, and a rendezvous and docking terminal; Obtaining a flag quantity for characterizing whether the tracking spacecraft is faulty, so as to determine a far guidance terminal from the co-orbital guidance terminal and the rendezvous and docking terminal; When the tracking spacecraft reaches the target anchor point, performing pulse guidance based on the position parameters of the target anchor point and the position parameters of the far guidance terminal, so that the tracking spacecraft reaches the far guidance terminal, and completing the long-range guidance of the tracking spacecraft; The position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal are determined by the following method: S1. Determining the spatial position of a target spacecraft that orbits and rendezvouses with the tracking spacecraft; S2. Based on the spatial position of the target spacecraft, determining a spatial position model of the target anchor point; S3. Based on the spatial position of the target spacecraft and the spatial position model of the target anchor point, determining the spatial position models of the co-orbital guidance terminal and the rendezvous and docking terminal; S4. Based on the spatial position models of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal, determining the initial values of the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal; S5. Based on the initial values of the position parameters, respectively determining the execution times of two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in two cases where the co-orbital guidance terminal is used as the far guidance terminal and the rendezvous and docking terminal is used as the far guidance terminal; S6. Respectively determining whether the execution times of the corresponding two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse meet the energy constraint and trajectory safety requirements in each case; S7. If at least one case does not meet the requirements, adjusting the initial values of the position parameters based on the spatial position models of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal, and jumping to execute S5 until the corresponding two guidance pulses in both cases meet the energy constraint and trajectory safety requirements, and obtaining the position parameters of the target anchor point, the co-orbital guidance terminal, and the rendezvous and docking terminal.

2. The method according to claim 1, characterized in that, The spatial position model of the target anchor point is: Among them, h p and h a satisfy the following formula: Where, θ0 is the phase angle difference between the target anchor point and the target spacecraft, and the target anchor point is located behind the target spacecraft, θ min and θ max are respectively the lower limit value and the upper limit value of the engineering constraint, x and z are respectively the relative positions of the rendezvous and docking terminal on the x-axis and the z-axis in the rendezvous and docking coordinate system, a T is the semi-major axis of the orbit of the target spacecraft, h p is the perigee altitude of the orbit where the target anchor point is located, h a is the apogee altitude of the orbit where the target anchor point is located, Δa p and Δa a are respectively the perigee orbit altitude difference and the apogee orbit altitude difference between the target anchor point and the rendezvous and docking terminal, θ1 is the phase angle interval of the tracking spacecraft from the target anchor point to the last guidance pulse before reaching the rendezvous and docking terminal, and θ2 is the phase angle interval of the tracking spacecraft from the last guidance pulse to the rendezvous and docking terminal.

3. The method according to claim 1, characterized in that, The spatial position model of the co-orbital guidance terminal is: The co-orbital guidance terminal is located behind the target spacecraft, and the phase angle difference between the co-orbital guidance terminal and the target spacecraft is θ s , and the altitude difference between the co-orbital guidance terminal and the target spacecraft is Δa; where θ s and Δa satisfy the following formula: Where, θ0 is the phase angle difference between the target anchor point and the target spacecraft, θ s is the phase angle difference between the co-orbital guidance terminal and the target spacecraft, θ1 is the phase angle interval of the tracking spacecraft from the target anchor point to the last guidance pulse implementation point before the rendezvous and docking terminal, Δa is the altitude difference between the co-orbital guidance terminal and the target spacecraft, a T is the semi-major axis of the target spacecraft's orbit, a C is the semi-major axis of the tracking spacecraft's orbit, and f(Δa) is the relationship function between the orbit altitude difference and the phase difference between the target spacecraft and the tracking spacecraft.

4. The method according to claim 1, characterized in that, The spatial position model of the rendezvous and docking terminal is: where ω T is the orbital angular velocity of the target spacecraft, x, y, and z are the relative positions of the rendezvous and docking terminal on the x-axis, y-axis, and z-axis respectively in the rendezvous and docking coordinate system, and v x , v x , and v x are the relative velocities of the rendezvous and docking terminal in the x-axis direction, y-axis direction, and z-axis direction respectively in the rendezvous and docking coordinate system.

5. The method according to claim 1, characterized in that, The respectively determining the execution times of two guidance pulses and the semi-major axis change amount corresponding to each guidance pulse in two cases where the co-orbital guidance terminal is used as the far guidance terminal and the rendezvous and docking terminal is used as the far guidance terminal based on the initial values of the position parameters includes: Taking the rendezvous and docking terminal as the far guidance terminal, and based on the initial value of the position parameter of the target anchor point, the initial value of the position parameter of the rendezvous and docking terminal, and the combined correction guidance algorithm, determining the latitude argument corresponding to the execution times of the two guidance pulses in this first case and the semi-major axis change amount corresponding to each guidance pulse; Taking the common-rail guidance terminal as the remote guidance terminal, based on the initial value of the position parameter of the target anchor point, the initial value of the position parameter of the common-rail guidance terminal, and the combined correction guidance algorithm, determine the latitude argument corresponding to the execution times of the two guidance pulses in this second case and the semi-major axis change amount corresponding to each of the guidance pulses.

6. The method according to any one of claims 1 - 5, characterized in that, The obtaining of the flag quantity used to characterize whether the tracking spacecraft is faulty to determine the remote guidance terminal from the common-rail guidance terminal and the rendezvous and docking terminal includes: When the flag quantity represents that the tracking spacecraft is faulty, determining the rendezvous and docking terminal as the remote guidance terminal; When the flag quantity represents that the tracking spacecraft is normal, determining the common-rail guidance terminal as the remote guidance terminal.

7. A remote guidance device based on anchor - point integrated common - rail and direct rendezvous, for implementing the method according to any one of claims 1 - 6, characterized in that, Applied to the guidance system of a tracking spacecraft, the device includes: An obtaining unit, configured to obtain the position parameters of a target anchor point, a common-rail guidance terminal, and a rendezvous and docking terminal; A determining unit, configured to obtain a flag quantity used to characterize whether the tracking spacecraft is faulty to determine the remote guidance terminal from the common-rail guidance terminal and the rendezvous and docking terminal; A guidance unit, configured to perform pulse guidance based on the position parameter of the target anchor point and the position parameter of the remote guidance terminal when the tracking spacecraft reaches the target anchor point, so that the tracking spacecraft reaches the remote guidance terminal to complete the remote guidance of the tracking spacecraft.

8. A computing device, comprising a memory and a processor, wherein a computer program is stored in the memory, and when the processor executes the computer program, the method according to any one of claims 1-6 is implemented.

9. A computer-readable storage medium, on which a computer program is stored, and when the computer program is executed in a computer, the computer is caused to execute the method according to any one of claims 1-6.

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