Rapid rendezvous flight method and device under solar angle energy constraints

Abstract

The present invention relates to the field of rendezvous and docking technology, and in particular to a method and device for rapid rendezvous flight under solar angle energy constraints. The method comprises: after a set time after a tracking spacecraft enters orbit, when the solar altitude angle is not within a solar altitude angle threshold interval, entering a continuous yaw flight mode; for each autonomous orbit control, executing: within a preset time interval before the start-up time of the autonomous orbit control, exiting the continuous yaw flight mode, and adjusting the flight attitude of the tracking spacecraft from the current yaw attitude to the orbit control pointing attitude of the autonomous orbit control; after the autonomous orbit control ends, adjusting the flight attitude of the tracking spacecraft from the orbit control pointing attitude corresponding to the autonomous orbit control to the target yaw attitude, and then switching to the continuous yaw flight mode; until the set conditions are met, switching from the continuous yaw flight mode to the target flight mode for rendezvous and docking. This scheme can greatly increase the launch time of the tracking spacecraft.

Description

Technical Field

[0001] The embodiments of the present invention relate to the technical field of rendezvous and docking, and in particular to a method and device for rapid rendezvous flight under solar angle energy constraints. Background Art

[0002] Because the angle between the sun and the orbital plane of the tracking spacecraft, that is, the solar altitude angle, is different at different times. Under the existing rapid rendezvous flight method, the tracking spacecraft is in a 0° attitude to the earth most of the time, and the solar sail panels usually rotate on a single axis. If the solar altitude angle of the orbital plane where the tracking spacecraft is located when entering orbit is large, the lateral incident angle of sunlight on the solar sail panels will be large. When it is large to a certain extent, the solar sail panels will have insufficient power generation efficiency and cannot meet the energy needs of the tracking spacecraft. Therefore, when selecting the launch window for the rendezvous and docking mission of the tracking spacecraft, it will be subject to the energy constraints of the solar altitude angle, resulting in the launch time of the tracking spacecraft being relatively limited, making it difficult to ensure the smooth implementation of various normal and emergency tasks during the space station operation phase.

[0003] Therefore, a new method of rapid rendezvous flight under solar angle energy constraints is urgently needed. Summary of the invention

[0004] In order to solve the problem that the launch window of the tracking spacecraft is constrained by the solar altitude angle energy in the traditional rapid rendezvous flight method, an embodiment of the present invention provides a rapid rendezvous flight method and device under the solar angle energy constraint.

[0005] In a first aspect, an embodiment of the present invention provides a method for rapid rendezvous flight under solar angle energy constraints, which is applied to a control system of a tracking spacecraft, and the method includes:

[0006] Obtaining a preset solar altitude angle threshold interval and a set time for enabling a continuous yaw flight mode after entering orbit; wherein the continuous yaw flight mode is used to adjust the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft satisfies both measurement and control conditions and energy conditions;

[0007] After the tracking spacecraft enters orbit, after the set time, determining whether the solar altitude angle at each time is within the solar altitude angle threshold range;

[0008] When the solar altitude angle is not within the solar altitude angle threshold range, entering a continuous yaw flight mode;

[0009] For each autonomous orbit control, the following operations are performed:

[0010] At a preset time interval before the start-up time of the autonomous orbit control, exit the continuous yaw flight mode, and adjust the flight attitude of the tracking spacecraft from the current yaw attitude to the orbit control pointing attitude corresponding to the autonomous orbit control;

[0011] When the autonomous orbit control is completed, the flight attitude of the tracking spacecraft is adjusted from the orbit control pointing attitude corresponding to the autonomous orbit control to the target yaw attitude, and then the tracking spacecraft enters the continuous yaw flight mode; the target yaw attitude is obtained based on the flight direction component and out-of-plane component of the current sun vector in the tracking spacecraft orbit coordinate system;

[0012] When the set conditions are met, the continuous yaw flight mode is switched to the target flight mode, so as to utilize the rendezvous sensor of the tracking spacecraft to rendezvous and dock with the target spacecraft; wherein, the target flight mode is a 0° attitude flight over the ground.

[0013] In a second aspect, an embodiment of the present invention further provides a rapid rendezvous flight device under solar angle energy constraints, which is arranged in a control system of a tracking spacecraft, and the device comprises:

[0014] an acquisition unit, used to acquire a preset solar altitude angle threshold interval and a set time for enabling a continuous yaw flight mode after entering orbit; wherein the continuous yaw flight mode is used to adjust the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft satisfies both measurement and control conditions and energy conditions;

[0015] A judging unit, configured to judge whether the solar altitude angle at each moment is within the solar altitude angle threshold interval after the tracking spacecraft enters orbit and after the set moment;

[0016] An opening unit, used for entering a continuous yaw flight mode when the solar altitude angle is not within the solar altitude angle threshold interval;

[0017] The orbit control preparation unit is used to execute, for each autonomous orbit control, the following steps: exiting the continuous yaw flight mode at a preset time interval before the start time of the autonomous orbit control, and adjusting the flight attitude of the tracking spacecraft from the current yaw attitude to the orbit control pointing attitude corresponding to the autonomous orbit control;

[0018] The orbit control ending unit is used for, after the autonomous orbit control ends, adjusting the flight attitude of the tracking spacecraft from the orbit control pointing attitude corresponding to the autonomous orbit control to the target yaw attitude, and then switching to a continuous yaw flight mode; the target yaw attitude is obtained based on the flight direction component and out-of-plane component of the current sun vector in the tracking spacecraft orbit coordinate system;

[0019] The rendezvous unit is used to switch from the continuous yaw flight mode to the target flight mode until the set conditions are met, so as to utilize the rendezvous sensor of the tracking spacecraft to rendezvous and dock with the target spacecraft; wherein, the target flight mode is a 0° attitude flight to the ground.

[0020] In a third aspect, an embodiment of the present invention further provides a computing device, including a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, the method described in any embodiment of this specification is implemented.

[0021] In a fourth aspect, an embodiment of the present invention further provides a computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, enables the computer to execute the method described in any embodiment of this specification.

[0022] The embodiment of the present invention provides a method and device for rapid rendezvous flight under solar angle energy constraints. First, a preset solar altitude angle threshold interval and a set time for allowing the continuous yaw flight mode to be turned on after entering orbit are obtained; wherein the continuous yaw flight mode is used to adjust the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft satisfies both the measurement and control conditions and the energy conditions; then, after the set time after the tracking spacecraft enters orbit, it is determined whether the solar altitude angle at each moment is within the solar altitude angle threshold interval; when the solar altitude angle is not within the solar altitude angle threshold interval, the continuous yaw flight mode is entered. ; and during each autonomous orbit control, the following are executed: at a preset time interval before the start-up time of this autonomous orbit control, the continuous yaw flight mode is exited, and the flight attitude of the tracking spacecraft is adjusted from the current yaw attitude to the orbit control pointing attitude corresponding to this autonomous orbit control; when this autonomous orbit control is completed, the flight attitude of the tracking spacecraft is adjusted from the orbit control pointing attitude corresponding to this autonomous orbit control to the target yaw attitude, and then the continuous yaw flight mode is entered; until the set conditions are met, the continuous yaw flight mode is switched to the target flight mode, so as to use the rendezvous sensor of the tracking spacecraft to rendezvous and dock with the target spacecraft. According to this scheme, after the set moment when the continuous yaw flight mode is allowed to be turned on after entering orbit, when the solar altitude angle is within the solar altitude angle threshold interval, the tracking spacecraft is made to fly at a 0° attitude relative to the earth. When the solar altitude angle exceeds the solar altitude angle threshold interval, it is switched to the continuous yaw flight mode. Only during the final rendezvous and docking and each autonomous orbit control, it is switched to the 0° attitude relative to the earth and the orbit control pointing attitude respectively. The continuous yaw flight mode is used at all other times. The continuous yaw flight mode adjusts the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft meets both the measurement and control conditions and the energy conditions, thereby eliminating the problem that the launch window of the tracking spacecraft is constrained by the solar altitude angle energy, thereby greatly increasing the launch time of the tracking spacecraft and enabling various normal and emergency tasks to be smoothly implemented during the space station operation phase. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0024] Figure 1 This is a flow chart of a rapid rendezvous flight method under solar angle energy constraints provided by one embodiment of the present invention;

[0025] Figure 2 is a hardware architecture diagram of a computing device provided by an embodiment of the present invention;

[0026] Figure 3 This is a structural diagram of a rapid rendezvous flight device under solar angle energy constraints provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0027] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0028] As mentioned above, the existing rapid rendezvous flight method will adjust the flight attitude of the tracking spacecraft to the orbit control pointing attitude only during each automatic orbit control, and the rest are all 0° attitude flights to the ground. This is because the 0° attitude flight to the ground can simultaneously meet the measurement and control needs, low energy consumption, low safety risks, simple flight mode and convenient flight mode switching. However, the solar sails of the tracking spacecraft in the low-Earth orbit are usually single-axis rotation. When the spacecraft is in orbit and flies at 0° attitude to the ground most of the time, when the sun is in the orbital plane of the tracking spacecraft, the solar sails can be rotated to face the sun; when there is an angle between the sun and the orbital plane of the tracking spacecraft (i.e., the solar altitude angle), even if the solar sails are controlled to face the sun, the sunlight is still incident from the side. The larger the solar altitude angle, the larger the angle of lateral incidence. When the lateral incidence angle is large to a certain extent, the solar sail power generation efficiency will be insufficient and it will not be able to meet the energy needs of the tracking spacecraft. Therefore, under this solar altitude angle energy constraint, when choosing the launch window for the rendezvous and docking mission of the tracking spacecraft, it is necessary to choose a launch time when the solar altitude angle is smaller after entering orbit. This will result in the launch time of the tracking spacecraft being relatively limited, making it difficult to ensure the smooth implementation of various normal and emergency tasks during the space station operation phase.

[0029] In order to solve the above technical problems, the inventor can consider that after the tracking spacecraft enters orbit, the continuous yaw flight mode is allowed to be turned on at the set time, and when the solar altitude angle is greater than the threshold range, the continuous yaw flight mode is entered. Except for the need to adjust to the orbit control pointing attitude before each autonomous orbit control and the need to adjust to the 0° attitude flight to the ground when the rendezvous sensor is needed for rendezvous and docking, the continuous yaw flight mode is used at other times; and the continuous yaw flight mode can adjust the flight attitude of the tracking spacecraft in real time, so that the flight attitude of the tracking spacecraft meets both the measurement and control conditions and the energy conditions. Then, even if the solar altitude angle is large, the tracking spacecraft can be launched, because the continuous yaw flight mode can be automatically switched to ensure energy requirements. Therefore, this scheme can eliminate the influence of the energy constraint of the solar angle on the launch window of the tracking spacecraft, and the launch time of the tracking spacecraft can be greatly increased, so that various normal and emergency tasks in the space station operation phase can be smoothly implemented.

[0030] The specific implementation of the above concept is described below.

[0031] Please refer to Figure 1 The embodiment of the present invention provides a rapid rendezvous flight method under solar angle energy constraints, which is applied to a control system of a tracking spacecraft. The method includes:

[0032] Step 100, obtaining a preset solar altitude angle threshold interval and a set time for enabling a continuous yaw flight mode after entering orbit; wherein the continuous yaw flight mode is used to adjust the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft satisfies both measurement and control conditions and energy conditions;

[0033] Step 102, after the tracking spacecraft enters orbit, after the set time, determine whether the solar altitude angle at each moment is within the solar altitude angle threshold range;

[0034] Step 104, when the solar altitude angle is not within the solar altitude angle threshold interval, entering the continuous yaw flight mode;

[0035] Step 106, for each autonomous orbit control, is performed: at a preset time interval before the start time of the autonomous orbit control, the continuous yaw flight mode is exited, and the flight attitude of the tracking spacecraft is adjusted from the current yaw attitude to the orbit control pointing attitude corresponding to the autonomous orbit control;

[0036] Step 108, after the autonomous orbit control is completed, the flight attitude of the tracking spacecraft is adjusted from the orbit control pointing attitude corresponding to the autonomous orbit control to the target yaw attitude, and then the continuous yaw flight mode is entered; the target yaw attitude is obtained based on the flight direction component and the out-of-plane component of the current sun vector in the tracking spacecraft orbit coordinate system;

[0037] Step 110, until the set conditions are met, the continuous yaw flight mode is switched to the target flight mode, so as to use the rendezvous sensor of the tracking spacecraft to rendezvous and dock with the target spacecraft; wherein the target flight mode is a 0° attitude flight over the ground.

[0038] In the embodiment of the present invention, first, a preset solar altitude angle threshold interval and a set time for allowing the continuous yaw flight mode to be turned on after entering orbit are obtained; wherein the continuous yaw flight mode is used to adjust the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft satisfies both the measurement and control conditions and the energy conditions; then, after the set time after the tracking spacecraft enters orbit, it is determined whether the solar altitude angle at each moment is within the solar altitude angle threshold interval; when the solar altitude angle is not within the solar altitude angle threshold interval, the continuous yaw flight mode is entered; and at each autonomous orbit control , all execute: at the preset time interval before the start-up time of this autonomous orbit control, exit the continuous yaw flight mode, and adjust the flight attitude of the tracking spacecraft from the current yaw attitude to the orbit control pointing attitude corresponding to this autonomous orbit control; when this autonomous orbit control ends, adjust the flight attitude of the tracking spacecraft from the orbit control pointing attitude corresponding to this autonomous orbit control to the target yaw attitude, and then enter the continuous yaw flight mode; until the set conditions are met, switch from the continuous yaw flight mode to the target flight mode, so as to use the rendezvous sensor of the tracking spacecraft to rendezvous and dock with the target spacecraft. According to this scheme, after the set moment when the continuous yaw flight mode is allowed to be turned on after entering orbit, when the solar altitude angle is within the solar altitude angle threshold interval, the tracking spacecraft is made to fly at a 0° attitude relative to the earth. When the solar altitude angle exceeds the solar altitude angle threshold interval, it is switched to the continuous yaw flight mode. It is only switched to the conventional corresponding attitude during the final rendezvous and docking and each autonomous orbit control. The continuous yaw flight mode is used at all other times. The continuous yaw flight mode adjusts the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft meets both the measurement and control conditions and the energy conditions, thereby eliminating the problem that the launch window of the tracking spacecraft is constrained by the solar altitude angle energy, thereby greatly increasing the launch time of the tracking spacecraft and enabling various normal and emergency tasks to be smoothly implemented during the space station operation phase.

[0039] For step 100:

[0040] In an embodiment of the present invention, it is necessary to determine the solar altitude angle threshold range for the tracking spacecraft to switch to the continuous yaw flight mode based on the solar panel power generation efficiency and the spacecraft energy consumption efficiency, so that the tracking spacecraft can monitor the solar altitude angle in real time and autonomously switch to the continuous yaw flight mode.

[0041] In some embodiments, the solar altitude angle threshold interval is calculated by the following formula:

[0042] [A1,A2]=f(P min )

[0043] in,

[0044]

[0045] Where [A1, A2] is the solar altitude angle threshold interval, A1 and A2 represent the minimum and maximum angles of the threshold interval, respectively. min is the minimum threshold of solar panel power generation efficiency, f() is the relationship function between solar panel power generation efficiency and solar altitude angle, is the energy consumption efficiency of the tracking spacecraft, and k is the ratio parameter between the solar panel power generation efficiency and the energy consumption efficiency of the tracking spacecraft.

[0046] In this embodiment, by setting the ratio parameter of the solar panel power generation efficiency and the energy consumption efficiency of the tracking spacecraft, k is greater than 1, to determine the minimum solar panel power generation efficiency (i.e., the minimum threshold of the solar panel power generation efficiency) at which the tracking spacecraft can achieve energy balance. Then, based on the relationship function between the solar panel power generation efficiency and the solar altitude angle, the solar altitude angle threshold range at the minimum solar panel power generation efficiency is determined.

[0047] In addition, after entering orbit, the tracking spacecraft will perform a series of key actions, and when performing key actions, it needs to maintain a good flight attitude (i.e., flying at a 0° attitude relative to the ground). Therefore, it is necessary to reserve enough flight time to perform key actions after entering orbit and determine the set time t0 at which the continuous yaw flight mode is allowed to be turned on after the spacecraft enters orbit.

[0048] In some embodiments, the set time is determined based on the time when the spacecraft enters orbit and the time required for the spacecraft to perform key actions after entering orbit; wherein the key actions include eliminating orbit errors, deploying solar panels, and confirming the status of the platform.

[0049] In this embodiment, the set time t0 is obtained by adding the time required for executing the key action after the tracking spacecraft enters orbit to the time when the tracking spacecraft enters orbit.

[0050] Regarding step 102:

[0051] After determining the solar altitude angle threshold interval and setting time t0 in step 100, binding can be performed before the tracking spacecraft is launched. When the tracking spacecraft enters orbit, the solar altitude angle at each moment is calculated in real time according to the navigation calculation results, and it is determined whether the solar altitude angle at each moment is within the solar altitude angle threshold interval, so as to monitor the occurrence of the situation where the solar altitude angle does not meet the energy balance requirements.

[0052] Regarding step 104:

[0053] In some embodiments, in the continuous yaw flight mode, the flight attitude of the tracking spacecraft satisfies the measurement and control conditions, which means that the roll angle and pitch angle of the tracking spacecraft are always 0; the flight attitude of the tracking spacecraft satisfies the energy conditions, which means that the rotation direction of the solar panel of the tracking spacecraft is perpendicular to the direction of the sun vector.

[0054] In this embodiment, the continuous yaw flight mode is used to adjust the flight attitude of the tracking spacecraft in real time. The flight attitude is (roll angle, pitch angle, yaw angle). The purpose of continuous yaw flight is to ensure that quadrant I of the spacecraft is always facing the ground, that is, the roll angle and the pitch angle are 0, which can meet the measurement and control conditions, and to make the rotation direction of the single-axis driven solar panel perpendicular to the direction of the solar vector, that is, by adjusting the yaw angle, the solar panel can be made perpendicular to the direction of the solar vector to meet the energy conditions.

[0055] In some embodiments, in the continuous yaw flight mode, the flight attitude of the tracking spacecraft satisfies the energy condition, which means that the yaw angle of the tracking spacecraft is controlled to always be the target yaw angle;

[0056] The target yaw angle is calculated by the following formula:

[0057] ψ t =atan2(Soy t ,Sox t )

[0058] In the formula, ψ t is the target yaw angle at each moment, t is the moment, Soy t is the flight direction component of the current sun vector in the tracking spacecraft orbit coordinate system, Sox t is the out-of-plane component of the current sun vector in the tracking spacecraft orbit coordinate system.

[0059] Since the sun's position changes over time, it is necessary to calculate the target yaw angle in real time based on the real-time sun vector in the flight direction component of the tracking spacecraft orbital coordinate system and the out-of-plane component of the tracking spacecraft orbital coordinate system, and adjust the flight attitude of the tracking spacecraft in real time to (0, 0, ψ t ).

[0060] Regarding step 106:

[0061] In some implementations, the preset time interval before each autonomous orbit control start-up time is determined by the following formula:

[0062]

[0063] Where Δt is the preset time interval before the autonomous orbit control is turned on, g is the magnification factor, and σ max is the maximum value of the yaw flight attitude range, is the minimum value of the orbit control pointing attitude range corresponding to this autonomous orbit control, and v is the attitude maneuvering speed of the tracking spacecraft.

[0064] In this embodiment, before each autonomous orbital control, since it is necessary to change from the current yaw attitude in the continuous yaw flight mode to the orbital control pointing attitude corresponding to each autonomous orbital control, and since the attitude maneuvering of the tracking spacecraft requires a certain amount of time, it is necessary to leave enough time for adjusting the attitude (i.e., the preset time interval) before each autonomous orbital control is started. After adjusting to the orbital control pointing attitude, the tracking spacecraft can perform autonomous orbital control.

[0065] Regarding step 108:

[0066] In the embodiment of the present invention, after each autonomous orbit control is performed, the flight attitude of the tracking spacecraft can be adjusted from the orbit control pointing attitude to the target yaw attitude, so as to enter the continuous yaw flight mode after the attitude adjustment.

[0067] It can be understood that the target yaw attitude is (0, 0, ψ t ), and the target yaw angle of the target yaw attitude is calculated by the following formula:

[0068] ψ t =atan2(Soy t ,Sox t )

[0069] In the formula, ψ t is the target yaw angle at each moment, t is the moment, Soy t is the flight direction component of the current sun vector in the tracking spacecraft orbit coordinate system, Sox t is the out-of-plane component of the current sun vector in the tracking spacecraft orbit coordinate system.

[0070] In addition, in a rendezvous and docking mission, multiple autonomous orbit controls may be performed. In this case, steps 106-108 need to be performed multiple times.

[0071] Regarding step 110:

[0072] In some embodiments, the setting conditions are: the energy balance maintenance time at this time is greater than the remaining mission flight time, and the distance between the tracking spacecraft and the target spacecraft at this time is less than the working distance of the rendezvous sensor of the tracking spacecraft.

[0073] In this embodiment, since the rendezvous sensor is usually configured at the head of the tracking spacecraft, when entering the working range of the relative measurement sensor (i.e., the rendezvous sensor), the head of the tracking spacecraft needs to be oriented to the target in order to obtain stable relative measurement, that is, it is necessary to maintain a 0° attitude flight relative to the ground (0, 0, 0). Therefore, it is necessary to determine whether the energy stored in the previous continuous yaw flight mode can enable the tracking spacecraft to maintain energy balance for a time sufficient to support subsequent tasks, and it is necessary to switch to the target flight mode only when the target tracking spacecraft enters the working range of the rendezvous sensor.

[0074] It can be understood that it is possible to determine in real time whether the two conditions of energy balance persistence time being greater than the remaining mission flight time and the distance between the tracking spacecraft and the target spacecraft being less than the working distance of the rendezvous sensor of the tracking spacecraft are simultaneously met, and when these two conditions are met, the continuous yaw flight mode is switched to the target flight mode. The time t for switching from the continuous yaw flight mode to the target flight mode can also be reasonably designed in advance based on comprehensive factors such as balanced energy, relative measurement and subsequent flight time. f , and t f Therefore, the setting condition can be that the energy balance persistence time at this time is greater than the remaining mission flight time, and the distance between the tracking spacecraft and the target spacecraft at this time is less than the working distance of the rendezvous sensor of the tracking spacecraft, or the time t f .

[0075] The embodiments of the present invention can relieve the launch window selection constraints to the greatest extent possible and realize a rapid autonomous rendezvous and docking mission under conditions of large solar altitude angles.

[0076] like Figure 2 , Figure 3 As shown, the embodiment of the present invention provides a fast rendezvous flight device under the constraint of solar angle energy. The device embodiment can be implemented by software, or by hardware or a combination of software and hardware. From the hardware level, Figure 2 As shown, it is a hardware architecture diagram of a computing device where a rapid rendezvous flight device under solar angle energy constraints is provided in an embodiment of the present invention. Figure 2 In addition to the processor, memory, network interface, and non-volatile memory shown in the figure, the computing device in which the device is located in the embodiment may also generally include other hardware, such as a forwarding chip responsible for processing messages, etc. Taking software implementation as an example, Figure 3 As shown, as a device in a logical sense, the CPU of the computing device in which it is located reads the corresponding computer program in the non-volatile memory into the memory and runs it. This embodiment provides a rapid rendezvous flight device under solar angle energy constraints, which is set in the control system of the tracking spacecraft, and the device includes:

[0077] The acquisition unit 301 is used to acquire a preset solar altitude angle threshold interval and a set time for enabling the continuous yaw flight mode after entering orbit; wherein the continuous yaw flight mode is used to adjust the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft satisfies both the measurement and control conditions and the energy conditions;

[0078] The judgment unit 302 is used to judge whether the solar altitude angle at each moment is within the solar altitude angle threshold range after the tracking spacecraft enters orbit and after the set time;

[0079] The start unit 303 is used to enter the continuous yaw flight mode when the solar altitude angle is not within the solar altitude angle threshold interval;

[0080] The orbit control preparation unit 304 is used to execute, for each autonomous orbit control, the following steps: exit the continuous yaw flight mode at a preset time interval before the start time of the autonomous orbit control, and adjust the flight attitude of the tracking spacecraft from the current yaw attitude to the orbit control pointing attitude corresponding to the autonomous orbit control;

[0081] The orbit control ending unit 305 is used to adjust the flight attitude of the tracking spacecraft from the orbit control pointing attitude corresponding to the autonomous orbit control to the target yaw attitude after the autonomous orbit control ends, and then switch to the continuous yaw flight mode; the target yaw attitude is obtained based on the flight direction component and out-of-plane component of the current sun vector in the orbit coordinate system of the tracking spacecraft;

[0082] The rendezvous unit 306 is used to switch from the continuous yaw flight mode to the target flight mode until the set conditions are met, so as to use the rendezvous sensor of the tracking spacecraft to rendezvous and dock with the target spacecraft; wherein the target flight mode is a 0° attitude flight over the ground.

[0083] In one embodiment of the present invention, the solar altitude angle threshold interval in the acquisition unit 301 is calculated by the following formula:

[0084] [A1,A2]=f(P min )

[0085] in,

[0086]

[0087] Where [A1, A2] is the solar altitude angle threshold interval, A1 and A2 represent the minimum and maximum angles of the threshold interval, respectively. min is the minimum threshold of solar panel power generation efficiency, f() is the relationship function between solar panel power generation efficiency and solar altitude angle, is the energy consumption efficiency of the tracking spacecraft, and k is the ratio parameter between the solar panel power generation efficiency and the energy consumption efficiency of the tracking spacecraft.

[0088] In one embodiment of the present invention, the set time in the acquisition unit 301 is determined based on the time when the tracking spacecraft enters orbit and the time required for the tracking spacecraft to perform key actions after entering orbit; wherein the key actions include eliminating orbit errors, unfolding solar panels, and confirming the platform status.

[0089] In one embodiment of the present invention, in the continuous yaw flight mode in the acquisition unit 301, the start unit 303, the orbit control preparation unit 304, and the orbit control end unit 305, the flight attitude of the tracking spacecraft satisfies the measurement and control conditions, which means that the roll angle and pitch angle of the tracking spacecraft are always 0; the flight attitude of the tracking spacecraft satisfies the energy conditions, which means that the rotation direction of the solar sail panel of the tracking spacecraft is perpendicular to the direction of the sun vector.

[0090] In one embodiment of the present invention, in the continuous yaw flight mode in the acquisition unit 301, the start unit 303, the orbit control preparation unit 304, and the orbit control end unit 305, the flight attitude of the tracking spacecraft meets the energy condition, which means that the yaw angle of the tracking spacecraft is always controlled to be the target yaw angle;

[0091] The target yaw angle is calculated by the following formula:

[0092] ψ t =atan2(Soy t ,Sox t )

[0093] In the formula, ψ t is the target yaw angle at each moment, t is the moment, Soy t is the flight direction component of the current sun vector in the tracking spacecraft orbit coordinate system, Sox t is the out-of-plane component of the current sun vector in the tracking spacecraft orbit coordinate system.

[0094] In one embodiment of the present invention, the preset time interval before each autonomous track control start-up time in the track control preparation unit 304 is determined by the following formula:

[0095]

[0096] Where Δt is the preset time interval before the autonomous orbit control is turned on, g is the magnification factor, and σ max is the maximum value of the yaw flight attitude range, is the minimum value of the orbit control pointing attitude range corresponding to this autonomous orbit control, and v is the attitude maneuvering speed of the tracking spacecraft.

[0097] In one embodiment of the present invention, the conditions set in the rendezvous unit 306 are: the energy balance maintenance time at this time is greater than the remaining mission flight time, and the distance between the tracking spacecraft and the target spacecraft at this time is less than the working distance of the rendezvous sensor of the tracking spacecraft.

[0098] It is to be understood that the structure illustrated in the embodiment of the present invention does not constitute a specific limitation on a fast rendezvous flight device under a solar angle energy constraint. In other embodiments of the present invention, a fast rendezvous flight device under a solar angle energy constraint may include more or fewer components than shown in the figure, or combine some components, or split some components, or arrange the components differently. The components shown in the figure may be implemented in hardware, software, or a combination of software and hardware.

[0099] The information interaction, execution process and other contents between the modules in the above-mentioned device are based on the same concept as the embodiment of the method of the present invention. For the specific contents, please refer to the description in the embodiment of the method of the present invention, and no further description is given here.

[0100] An embodiment of the present invention further provides a computing device, including a memory and a processor, wherein a computer program is stored in the memory, and when the processor executes the computer program, a rapid rendezvous flight method under solar angle energy constraints in any embodiment of the present invention is implemented.

[0101] An embodiment of the present invention further provides a computer-readable storage medium having a computer program stored thereon. When the computer program is executed by a processor, the processor executes a rapid rendezvous flight method under solar angle energy constraints in any embodiment of the present invention.

[0102] Specifically, a system or device equipped with a storage medium can be provided, on which software program code that implements the functions of any of the above-mentioned embodiments is stored, and a computer (or CPU or MPU) of the system or device can read and execute the program code stored in the storage medium.

[0103] In this case, the program code itself read from the storage medium can realize the function of any one of the above-mentioned embodiments, and thus the program code and the storage medium storing the program code constitute a part of the present invention.

[0104] The storage medium embodiments for providing the program code include a floppy disk, a hard disk, a magneto-optical disk, an optical disk (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), a magnetic tape, a non-volatile memory card, and a ROM. Alternatively, the program code can be downloaded from a server computer by a communication network.

[0105] In addition, it should be clear that the functions of any of the above embodiments can be implemented not only by executing the program code read by the computer, but also by enabling an operating system operating on the computer to complete part or all of the actual operations based on instructions from the program code.

[0106] In addition, it can be understood that the program code read from the storage medium is written to a memory provided in an expansion board inserted into the computer or to a memory provided in an expansion module connected to the computer, and then based on the instructions of the program code, a CPU installed on the expansion board or expansion module is enabled to perform part or all of the actual operations, thereby realizing the functions of any of the above-mentioned embodiments.

[0107] 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 such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device.

[0108] A person 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, and the aforementioned program can be stored in a computer-readable storage medium. When the program is executed, it executes the steps of the above method embodiments; and the aforementioned storage medium includes: ROM, RAM, magnetic disk or optical disk, etc., various media that can store program codes.

[0109] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A rapid rendezvous flight method under solar angle energy constraints, characterized in that: A control system applied to a tracking spacecraft, the method comprising: Obtaining a preset solar altitude angle threshold interval and a set time for enabling a continuous yaw flight mode after entering orbit; wherein the continuous yaw flight mode is used to adjust the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft satisfies both measurement and control conditions and energy conditions; After the tracking spacecraft enters orbit, after the set time, determining whether the solar altitude angle at each time is within the solar altitude angle threshold range; When the solar altitude angle is not within the solar altitude angle threshold range, entering a continuous yaw flight mode; For each autonomous orbit control, the following operations are performed: At a preset time interval before the start-up time of the autonomous orbit control, exit the continuous yaw flight mode, and adjust the flight attitude of the tracking spacecraft from the current yaw attitude to the orbit control pointing attitude corresponding to the autonomous orbit control; When the autonomous orbit control is completed, the flight attitude of the tracking spacecraft is adjusted from the orbit control pointing attitude corresponding to the autonomous orbit control to the target yaw attitude, and then the tracking spacecraft enters the continuous yaw flight mode; the target yaw attitude is obtained based on the flight direction component and out-of-plane component of the current sun vector in the tracking spacecraft orbit coordinate system; When the set conditions are met, the continuous yaw flight mode is switched to the target flight mode, so as to utilize the rendezvous sensor of the tracking spacecraft to rendezvous and dock with the target spacecraft; wherein, the target flight mode is a 0° attitude flight over the ground.

2. The method according to claim 1, characterized in that: The solar altitude angle threshold interval is calculated by the following formula: [A1,A2]=f(P min ) in, Where [A1, A2] is the solar altitude angle threshold interval, A1 and A2 represent the minimum and maximum angles of the threshold interval, respectively. min is the minimum threshold of solar panel power generation efficiency, f() is the relationship function between solar panel power generation efficiency and solar altitude angle, is the energy consumption efficiency of the tracking spacecraft, and k is the ratio parameter of the solar panel power generation efficiency to the energy consumption efficiency of the tracking spacecraft.

3. The method according to claim 1, characterized in that The set time is determined based on the time when the tracking spacecraft enters orbit and the time required for the tracking spacecraft to perform key actions after entering orbit; wherein the key actions include eliminating orbit entry errors, unfolding solar panels, and confirming the platform status.

4. The method according to claim 1, characterized in that: In the continuous yaw flight mode, the flight attitude of the tracking spacecraft satisfies the measurement and control conditions, which means that the roll angle and pitch angle of the tracking spacecraft are always 0; the flight attitude of the tracking spacecraft satisfies the energy conditions, which means that the rotation direction of the solar panel of the tracking spacecraft is perpendicular to the direction of the sun vector, so that the solar panel is facing the sun at any solar altitude angle.

5. The method according to claim 4, characterized in that In the continuous yaw flight mode, the flight attitude of the tracking spacecraft satisfies the energy condition, which means that the yaw angle of the tracking spacecraft is always controlled to be the target yaw angle; The target yaw angle is calculated by the following formula: ψ t =atan2(Soy t ,Sox t ) In the formula, ψ t is the target yaw angle at each moment, t is the moment, Soy t is the flight direction component of the current sun vector in the tracking spacecraft orbit coordinate system, Sox t is the out-of-plane component of the current sun vector in the tracking spacecraft orbit coordinate system.

6. The method according to claim 1, characterized in that The preset time interval before each autonomous orbit control start-up is determined by the following formula: Where Δt is the preset time interval before the autonomous orbit control is turned on, g is the magnification factor, and σ max is the maximum value of the yaw flight attitude range, is the minimum value of the orbit control pointing attitude range corresponding to this autonomous orbit control, and v is the attitude maneuvering speed of the tracking spacecraft.

7. The method according to any one of claims 1 to 6, characterized in that The setting condition is: at this time, the energy balance maintenance time is greater than the remaining mission flight time, and at this time, the distance between the tracking spacecraft and the target spacecraft is less than the working distance of the rendezvous sensor of the tracking spacecraft.

8. A rapid rendezvous flight device under solar angle energy constraints, characterized in that: A control system provided for tracking a spacecraft, the device comprising: an acquisition unit, used to acquire a preset solar altitude angle threshold interval and a set time for enabling a continuous yaw flight mode after entering orbit; wherein the continuous yaw flight mode is used to adjust the flight attitude of the tracking spacecraft in real time so that the flight attitude of the tracking spacecraft satisfies both measurement and control conditions and energy conditions; A judging unit, configured to judge whether the solar altitude angle at each moment is within the solar altitude angle threshold range after the tracking spacecraft enters orbit and after the set moment; An opening unit, used for entering a continuous yaw flight mode when the solar altitude angle is not within the solar altitude angle threshold interval; The orbit control preparation unit is used to execute, for each autonomous orbit control, the following steps: exiting the continuous yaw flight mode at a preset time interval before the start time of the autonomous orbit control, and adjusting the flight attitude of the tracking spacecraft from the current yaw attitude to the orbit control pointing attitude corresponding to the autonomous orbit control; The orbit control ending unit is used for, after the autonomous orbit control ends, adjusting the flight attitude of the tracking spacecraft from the orbit control pointing attitude corresponding to the autonomous orbit control to the target yaw attitude, and then switching to a continuous yaw flight mode; the target yaw attitude is obtained based on the flight direction component and out-of-plane component of the current sun vector in the tracking spacecraft orbit coordinate system; The rendezvous unit is used to switch from the continuous yaw flight mode to the target flight mode until the set conditions are met, so as to utilize the rendezvous sensor of the tracking spacecraft to rendezvous and dock with the target spacecraft; wherein, the target flight mode is a 0° attitude flight to the ground.

9. A computing device, comprising a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, the method according to any one of claims 1 to 7 is implemented.

10. A computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to execute the method according to any one of claims 1 to 7.

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