Satellite attitude control method and system for satellite-ground alignment

Abstract

The present invention provides a satellite attitude control method and system for satellite-ground alignment. The method includes: obtaining output information from a payload detector, the output information including the coordinates of a light spot on the detector from a landmark laser through an optical lens; calculating the optical axis deviation angle between the payload reference optical axis and the landmark laser based on the output information; calculating the final satellite station vector based on the optical axis deviation angle, the final satellite station vector being a unit vector pointing from the satellite's center of mass to the ground communication station in the satellite's coordinate system; calculating the measured station angle based on the final satellite station vector; obtaining the satellite platform's attitude angle and angular velocity; calculating the satellite's three-axis control torque based on the attitude angle, the satellite platform's attitude angle, and the satellite platform's attitude velocity; and controlling the satellite platform's motion based on the satellite's three-axis control torque. The present invention utilizes the payload detector's output information as input information for the satellite platform's control system, eliminating the need for a coarse tracking system and simplifying the system.

Description

Technical Field

[0001] The present invention mainly relates to the field of satellite technology, and in particular to a satellite attitude control method and system for satellite-ground alignment. Background Art

[0002] When conducting space-to-ground quantum communication experiments, quantum science experimental satellites need to align the optical axis of the effective quantum payload on board with the ground communication station, and have very high requirements for the accuracy of space-to-ground alignment, up to 3 microradians (urad).

[0003] To address this challenge, a three-level control approach is currently being adopted, involving the satellite platform and payload's coarse and fine tracking systems, to ensure alignment accuracy. First, the satellite platform's three-axis attitude is controlled to ensure it points toward the Earth or toward a ground optical station. Next, the payload utilizes its own coarse tracking system to scan and capture a wide range of beacon light and guide it into the fine tracking field of view. The fine tracking system further reduces tracking errors based on the coarse tracking accuracy, ultimately achieving the required alignment accuracy. This entire control process is complex and time-consuming. Summary of the Invention

[0004] The technical problem to be solved by the present invention is to provide a satellite attitude control method and system for satellite-ground alignment, so as to solve the problem that the existing satellite-ground alignment control system is complex and time-consuming.

[0005] To solve the above technical problems, the present invention provides a satellite attitude control method for satellite-ground alignment, comprising: obtaining output information of a payload detector, the output information including the coordinates of a light spot of a landmark laser on the detector through an optical lens; calculating an optical axis deviation angle between a payload reference optical axis and a landmark laser according to the output information; calculating a final satellite station vector according to the optical axis deviation angle, the final satellite station vector being a unit vector pointing from the satellite center of mass to a ground communication station in a satellite body coordinate system; calculating a measured station angle according to the final satellite station vector; obtaining a satellite platform attitude angle and a satellite platform attitude angular velocity; calculating a satellite three-axis control torque according to the station angle, the satellite platform attitude angle and the satellite platform attitude angular velocity, and controlling the satellite platform motion according to the satellite three-axis control torque.

[0006] Optionally, the optical axis deviation angle is calculated using the following formula:

[0007] α=K cam (x-x0)

[0008] β=K cam (y-y0)

[0009] Wherein, α and β are the optical axis deviation angles, K camare constant coefficients related to the geometric distance and focal length of the payload detector pixels, x0 and y0 are the offset correction values ​​of the payload detector center point, and the coordinates of the light spot are (x, y, 0).

[0010] Optionally, calculating the final star station vector according to the optical axis deflection angle includes: calculating an initial star station vector according to the optical axis deflection angle, the initial star station vector being a vector pointing from the satellite center of mass to the ground communication station in the detector coordinate system; converting the initial star station vector into an intermediate star station vector according to a transformation matrix from the detector coordinate system to the satellite body coordinate system; and regularizing the intermediate star station vector to obtain the final star station vector.

[0011] Optionally, the initial station vector is calculated using the following formula:

[0012] ST m =[-α-β1] T

[0013] Wherein, α and β are the optical axis deviation angles.

[0014] Optionally, the measured subtended station angle is calculated using the following formula:

[0015]

[0016] θ m =-atan2(μST bx ,μST bz )

[0017]

[0018] Among them, Mesure_A bo is the measured station angle, and the final star station vector is (μST bx , μST by , μST bz ), the function f=atan2(m,n) is used to find the variable The inverse tangent value of , and the range is f∈-π,π.

[0019] Optionally, calculating the satellite three-axis control torque based on the station angle, the satellite platform attitude angle and the satellite platform attitude angular velocity includes: calculating the station error angle based on the station angle and a preset expected station angle; calculating the station axial error attitude angle based on the satellite platform attitude angle and the preset expected attitude angle; calculating the error angular velocity based on the satellite platform attitude angular velocity and the preset expected angular velocity; calculating the satellite three-axis control torque based on the station error angle, the station axial error attitude angle and the error angular velocity.

[0020] Optionally, the satellite three-axis control torque is calculated using the following formula:

[0021]

[0022]

[0023] Among them, K p =[k px k py k pz ]、K d =[k dx k dy k dz ]、K i =[k ix k iy k iz ] are the designed proportional, differential and integral control parameters respectively, T i_pre =[T ix_pre T iy_pre T iz_pre ] T T of the previous control cycle i , the initial value is T i0 =[00 0] T , station error angle and θ e , the station axial error attitude angle ψ e , error angular velocity ω ex 、ω ey 、ω ez .

[0024] Optionally, the satellite attitude control method further includes: obtaining a safe operating range of the satellite angular velocity, the safe operating range including a maximum forward angular velocity and a maximum reverse angular velocity; and limiting the station error angle and the station axial error attitude angle according to the safe operating range.

[0025] In order to solve the above technical problems, the present invention also provides a satellite attitude control system, including a precise tracking system for a payload, wherein the satellite attitude control system is used to align the payload reference optical axis with the ground communication station, and further includes: a detector of the payload, used to collect the coordinates of the light spot on the detector through the optical lens of the landmark laser, and send the coordinates of the light spot to the platform control system of the satellite platform; the satellite platform includes: an attitude sensor, used to measure the satellite platform attitude angle and the satellite platform attitude angular velocity; and a platform control system, used to execute the satellite attitude control method as described above.

[0026] Optionally, the satellite platform further includes: a storage unit which stores a safe operating range of the angular velocity of the satellite body; the platform control system is further used to: obtain the safe operating range, and limit the station error angle and the station axial error attitude angle according to the safe operating range.

[0027] Compared with the prior art, the present invention has the following advantages:

[0028] The satellite attitude control method and system of the present invention utilize the output information of the payload's detector as the input information of the satellite platform's control system. In this way, only one axis of the satellite platform needs to be controlled according to the installation method of the payload, thereby improving the satellite platform's pointing accuracy to the station, and then directly introducing the beacon light into the fine tracking field of view, thereby eliminating the coarse tracking system and greatly simplifying the satellite system. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] The accompanying drawings are included to provide a further understanding of the present application, are incorporated into and constitute a part of this application, illustrate embodiments of the present application, and together with this specification serve to explain the principles of the present invention. In the accompanying drawings:

[0030] Figure 1A This is a classic satellite system block diagram.

[0031] Figure 1B This is a schematic diagram of the traditional three-level control satellite platform and coarse tracking system.

[0032] Figure 2 4 is a system block diagram of a satellite attitude control system according to an embodiment of the present invention.

[0033] Figure 3 FIG. 1 is a schematic diagram of the deviation angle between the payload reference optical axis and the optical axis of the landmark laser according to an embodiment of the present invention.

[0034] Figure 4 is a flowchart of a satellite attitude control method according to an embodiment of the present invention.

[0035] Figure 5 FIG. 4 is a schematic diagram of a final satellite station vector according to an embodiment of the present invention.

[0036] Figure 6 yes Figure 4 Flowchart of an embodiment of step S43 in FIG.

[0037] Figure 7 yes Figure 4 Flowchart of an embodiment of step S46 in FIG.

[0038] Figure 8 yes Figure 4 Flowchart of another embodiment of step S46. DETAILED DESCRIPTION

[0039] To more clearly illustrate the technical solutions of the embodiments of this application, the following is a brief introduction to the drawings required for describing the embodiments. Obviously, the drawings described below are merely examples or embodiments of this application. Those skilled in the art can apply this application to other similar scenarios based on these drawings without inventive effort. Unless otherwise apparent from the context or otherwise noted, the same reference numerals in the figures represent the same structure or operation.

[0040] Figure 1A This is a classic satellite system block diagram. As shown in Figure 1, satellite 100 includes a satellite platform 11 and a payload 12. Payload 12 refers to the equipment, instruments or systems on the satellite that directly complete specific flight missions and is the core part of the satellite. For example, payloads include communications, navigation and positioning, earth observation, and space science and technology experiments. In this embodiment, the payload is used to complete quantum science experiments. Payload 12 includes a detector 121, a coarse tracking system 122, and a fine tracking system 123. Satellite platform 11 is a service system for payload 12, used to ensure the normal operation of all subsystems on the satellite from the takeoff of the carrier rocket to the end of its working life. Satellite platform 11 includes an attitude sensor 111 and a platform control system 112.

[0041] When conducting space-to-ground quantum communication experiments, quantum science experimental satellites must establish and maintain stable satellite-to-ground beacon and communication optical links between the quantum satellite and the ground station. To address this challenge, a three-level composite control approach is currently being adopted, involving a coarse tracking system for the satellite platform and payload, and a fine tracking system. Figure 1B This is a schematic diagram of the traditional three-level control satellite platform and coarse tracking system. Figure 1B As shown, when passing a station, the satellite platform first controls its three-axis attitude to point toward a ground communication station or a ground optical station. Based on this, the payload uses its own coarse tracking system to scan and capture a wide range of beacon light and guide the beacon light into the fine tracking field of view. The fine tracking system (not shown) further reduces tracking error based on the coarse tracking accuracy, ultimately achieving the required alignment accuracy. The control methods for achieving satellite-to-ground alignment in the existing technology are very complex. This is because the pointing process of a low-orbit satellite platform toward a ground optical station is a dynamic tracking process with large-angle rapid maneuvers, which places high demands on the measurement accuracy of the satellite platform's attitude sensor, as well as the dynamic response speed and tracking accuracy of the control system. At high tracking speeds, the measurement accuracy of the attitude sensor decreases, resulting in poor station pointing accuracy, currently only 0.5°, and inability to directly guide the beacon light into the payload's fine tracking field of view. The payload's coarse tracking system is required to further control the beacon light and bring it into the fine tracking field of view.

[0042] In view of the shortcomings of the above-mentioned prior art, the present invention aims to provide a satellite attitude control method and system based on satellite-station vector measurement information. This method only requires controlling a single axis of the satellite platform based on the payload's mounting configuration. This allows the output information from the payload's detectors to be used as input for the satellite platform's control system. This improves measurement accuracy and, due to the improved pointing accuracy of the satellite platform to the station, allows beacon light to be directly introduced into the fine tracking field of view, eliminating the need for a coarse tracking system and significantly simplifying the satellite system.

[0043] Figure 2 FIG1 is a system block diagram of a satellite attitude control system according to an embodiment of the present invention. The satellite attitude control system is used to align the payload reference optical axis with the ground communication station. Figure 2 As shown, satellite attitude control system 200 includes satellite platform 21 and payload 22. Satellite platform 21 includes attitude sensor 211 and platform control system 212. Payload 22 includes detector 221 and fine tracking system 222. Detector 221 is used to collect the coordinates of the light spot of the landmark laser on detector 221 through the optical lens and transmit the light spot coordinates to platform control system 212. Platform control system 212 is used to calculate the deviation angle between the payload reference optical axis and the optical axis of the landmark laser based on the light spot coordinates. Based on the optical axis deviation angle, the final satellite station vector is calculated based on the final satellite station vector. The final satellite station vector is a unit vector pointing from the satellite's center of mass to the ground communication station in the satellite's coordinate system. The final satellite station vector is used to calculate the measured station angle. Attitude sensor 211 is used to collect the satellite platform's attitude angle and angular velocity for attitude determination. Platform control system 212 is also used to calculate the satellite's three-axis control torque based on the station angle, the satellite platform's attitude angle, and the satellite platform's attitude angular velocity for attitude determination. The satellite platform's motion is controlled based on the satellite's three-axis control torque. As the satellite platform's pointing accuracy to the station is improved, the beacon light can be directly introduced into the fine tracking field of view. The fine tracking system 222 is used to scan and capture the beacon light and further reduce the tracking error based on the beacon light.

[0044] Figure 3 Schematic diagram of the deviation angle between the payload reference optical axis and the landmark laser optical axis according to an embodiment of the present invention. Figure 3 As shown, the detector coordinate system OXYZ is established with the center point of the payload detector as the coordinate origin O. The coordinates of the light spot projected by the landmark laser through the lens center O'(0, 0, H) of the optical lens on the end face of the detector are A(x, y, 0). The detector collects the x and y coordinates in real time and sends the x and y coordinates to the platform control system. B and C are the mapping of A on the X and Y axes respectively. The angle between the vector from B to the lens center O' and the Z axis is α, and the angle between the vector from C to O' and the Z axis is β. Among them, α and β are the optical axis deviation angles between the payload reference optical axis and the landmark laser.

[0045] Figure 4FIG. 1 is a flow chart of a satellite attitude control method according to an embodiment of the present invention. The satellite attitude control method is executed by a platform control system. Figure 4 As shown, the satellite attitude control method 400 includes:

[0046] Step S41: Obtain output information of the payload detector, where the output information includes the coordinates of the light spot of the landmark laser on the detector through the optical lens.

[0047] Step S42: Calculate the optical axis deviation angle between the payload reference optical axis and the landmark laser according to the output information.

[0048] Alternatively, the optical axis deviation angles α and β can be calculated using the following formula:

[0049] α=K cam (x-x0)

[0050] β=K cam (y-y0)

[0051] Among them, the coordinates of the light spot are (x, y, 0), K cam is a constant coefficient related to the geometric distance and focal length of the payload detector pixel, x0 and y0 are the correction values ​​of the payload detector center point offset, and K cam , x0, and y0 are obtained through ground tests and calibration after entering orbit.

[0052] Step S43: Calculate the final star station vector according to the optical axis deflection angle. Figure 5 is a schematic diagram of the final satellite station vector according to an embodiment of the present invention. Figure 5 As shown, the final star station vector μST b It is the unit vector pointing from the satellite's center of mass to the ground communication station in the satellite's body coordinate system.

[0053] Figure 6 yes Figure 4 Flowchart of step S43 in the embodiment. Figure 6 As shown, the calculation of the final star station vector based on the optical axis deflection angle includes:

[0054] Step S431: Calculate the initial satellite station vector according to the optical axis deflection. The initial satellite station vector is the vector pointing from the satellite's centroid to the ground communication station in the detector coordinate system.

[0055] Alternatively, the initial star station vector ST can be calculated using the following formula: m :

[0056] ST m =[-α -β 1] T

[0057] Among them, α and β are the optical axis deviation angles.

[0058] Step S432: Convert the initial star station vector into an intermediate star station vector according to the transformation matrix from the detector coordinate system to the satellite body coordinate system.

[0059] Alternatively, the intermediate station vector ST can be calculated using the following formula: b :

[0060] ST b =R bm_Cam ST m

[0061] where R bm_Cam It is the transformation matrix from the payload's detector coordinate system to the satellite's body coordinate system, which is determined by the installation relationship of the payload relative to the satellite.

[0062] Step S433: Regularize the intermediate satellite station vector to obtain the final satellite station vector.

[0063] Alternatively, the final station vector μST can be calculated using the following formula: b :

[0064]

[0065] where ||ST b || Vector ST b The modulus value of .

[0066] Step S44: Calculate the measured station angle Mesure_A based on the final satellite station vector bo .

[0067] Taking the satellite + Z axis alignment as an example, the alignment angle Mesure-A can be calculated using the following formula: bo :

[0068]

[0069] θ m =-atan2(μSTb x , μST bz )

[0070]

[0071] Among them, the final star station vector is (μST bx , μST by , μST bz ), the function f=atan2(m,n) is used to find the variable The inverse tangent value of , and the range is f∈[-π,π].

[0072] Although satellite + Z-axis alignment is usually used in actual engineering, it should be noted that the present invention is not limited to satellite + Z-axis alignment. In actual use, the calculation formula in the specific implementation method can be adaptively modified according to the different axial directions of the satellite alignment.

[0073] Step S45: Acquire the satellite platform attitude angle and the satellite platform attitude angular velocity.

[0074] The platform control system 212 can obtain the satellite platform attitude angle from the attitude sensor 211 and the satellite platform attitude angular velocity Attω bo =[ω box ω boy ω boz ].

[0075] Step S46: Calculate the satellite three-axis control torque based on the station angle, the satellite platform attitude angle and the satellite platform attitude angular velocity, and control the satellite platform movement based on the satellite three-axis control torque.

[0076] Figure 7 yes Figure 4 Flowchart of step S46 in the embodiment. Figure 7 As shown in FIG, the satellite three-axis control torque is calculated based on the station angle, the satellite platform attitude angle and the satellite platform attitude angular velocity, including:

[0077] Step S461: Calculate the alignment error angle based on the alignment angle and the preset expected alignment angle.

[0078] Taking the preset expected station angle as [0 0 0] as an example, the measured station angle Calculate the station error angle by adding the expected station angle [0 0 0] and θ e , which serves as the input for the subsequent calculation of the control torque required for station pointing.

[0079]

[0080] θ e =θ m

[0081] Step S462: Determine the attitude angle of the satellite platform and the preset desired attitude angle Calculate the station axial error attitude angle ψ e .

[0082] Optionally, the station axial error attitude angle ψ is calculated using the following formula: e , as the input for the subsequent calculation of the control torque required to stably maintain the satellite's axial attitude to the station to ensure the consistency of the polarization state of quantum communication:

[0083]

[0084]

[0085]

[0086] ψ e =atan2(-R bf (2,1),R bf (2,2))

[0087] Among them, R fo is the transformation matrix from the orbital coordinate system to the reference coordinate system, R bo is the transformation matrix from the orbital coordinate system to the satellite body coordinate system, is the matrix R fo The transposed matrix, R bf (2,1) is the matrix R bf The element in row 2 and column 1, R bf (2,2) is the matrix R bf The element at row 2 and column 2.

[0088] Step S463: Determine the attitude angular velocity Attω of the satellite platform bo =[ω box ω boy ω boz ] and the preset desired angular velocity ω boc =[ω bocx ω bocy ω bocz ]Calculate the error angular velocity ω ex 、ω ey 、ω ez .

[0089] Alternatively, the error angular velocity ω is calculated by the following formula: ex 、ω ey 、ω ez :

[0090]

[0091]

[0092]

[0093] Among them, R bf (3,:)=[R bf (3,1) R bf (3,2) R bf (3,2)], is the vector ω boc The transpose of Rbf (3,1) is the matrix R bf The element in row 3 and column 1, R bf (3,2) is the matrix R bf The element in row 3 and column 2, R bf (3,3) is the matrix R bf The element at row 3 and column 3.

[0094] Step S464: Based on the station error angle and θ e 、Station axial error attitude angle ψ e and error angular velocity ω ex 、ω ey 、ω ez Calculate the satellite's three-axis control torque T c .

[0095] Optionally, the satellite three-axis control torque T is calculated by the following formula: c :

[0096]

[0097]

[0098] Among them, K p =[k px k py k pz ]、K d =[k dx k dy k dz ]、K i =[k ix k iy k iz ] are the designed proportional, differential and integral control parameters respectively, T i_pre =[T ix_pre T iy_pre T iz_pre ] T T of the previous control cycle i , the initial value is T i0 =[00 0] T .

[0099] The satellite attitude control method for satellite-ground alignment of the present invention fully utilizes the high-precision information output by the payload at a high frequency to calculate the final satellite station vector. The final satellite station vector is directly used as the control input, and its component perpendicular to the target station axis is controlled to converge to 0, thereby achieving ultra-high-precision station-pointing attitude control. At the same time, in order to meet the polarization state consistency requirements of quantum communication, the satellite attitude in the target station axis is stably maintained and controlled to maintain the desired attitude agreed with the ground communication station.

[0100] Figure 8 yes Figure 4 Flowchart of another embodiment of step S46. Figure 8 As shown, the calculation of the satellite three-axis control torque based on the station angle, the satellite platform attitude angle and the satellite platform attitude angular velocity also includes:

[0101] Step S465: Obtain the safe operating range of the angular velocity of the star, where the safe operating range includes the maximum forward angular velocity and the maximum reverse angular velocity.

[0102] Step S466: limiting the station error angle and the station axial error attitude angle according to the safe operating range.

[0103] During engineering use, for the sake of the safety of the entire satellite, the angular velocity of the satellite is limited to [-ω max ,ω max ], where [-ω max ,ω max ] is the interval from the maximum value of the reverse angular velocity to the maximum value of the forward angular velocity, and the station error angle and θ e 、Station axial error attitude angle ψ e To perform limiting processing:

[0104] like but Otherwise if but otherwise No action is taken.

[0105] If θ e ≥k dy ω max / k py , then θ e =k dy ω max / k py Otherwise, if θ e ≤-k dy ω max / k py , then θ e =-k dy ω max / k py ; otherwise θ e No action is taken.

[0106] If ψ e ≥k dz ω max / k pz , then ψ e =k dz ω max / kpz ; Otherwise, if ψ e ≤-k dz ω max / k pz , then ψ e =-k dz ω max / k pz ; otherwise ψ e No action is taken.

[0107] where K p =[k px k py k pz ]、K d =[k dx k dy k dz ] are the designed proportional and differential control parameters respectively.

[0108] Correspondingly, in some embodiments, the satellite platform also includes: a storage unit, which stores the safe operating range of the satellite angular velocity. The platform control system 212 is also used to obtain the safe operating range from the storage unit, and limit the station error angle and the station axial error attitude angle according to the safe operating range.

[0109] The basic concepts have been described above. It will be apparent to those skilled in the art that the above disclosures are merely illustrative and do not constitute limitations on this application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements, and amendments to this application. Such modifications, improvements, and amendments are suggested in this application and remain within the spirit and scope of the exemplary embodiments of this application.

[0110] As used in this application and the claims, unless the context clearly indicates otherwise, the words "a," "an," "an," and / or "the" are not intended to refer to the singular but may include the plural. Generally speaking, the terms "comprises" and "include" only indicate the inclusion of the steps and elements specifically identified, and these steps and elements do not constitute an exclusive list. A method or apparatus may also include other steps or elements.

[0111] Unless otherwise specifically stated, the relative arrangement of the parts and steps, numerical expressions and numerical values ​​set forth in these embodiments do not limit the scope of the present application. At the same time, it should be understood that, for ease of description, the sizes of the various parts shown in the drawings are not drawn according to actual proportional relationships. The techniques, methods and equipment known to those of ordinary skill in the relevant art may not be discussed in detail, but where appropriate, the techniques, methods and equipment should be considered as part of the authorization specification. In all examples shown and discussed here, any specific values ​​should be interpreted as being merely exemplary, not as limitations. Therefore, other examples of the exemplary embodiments may have different values. It should be noted that similar numbers and letters represent similar items in the following figures, and therefore, once an item is defined in one figure, it does not need to be further discussed in subsequent figures.

[0112] Flowcharts are used in this application to illustrate the operations performed by systems according to embodiments of the present application. It should be understood that the preceding or following operations are not necessarily performed in exact order. Instead, the various steps may be processed in reverse order or simultaneously. Furthermore, other operations may be added to these processes, or one or more operations may be removed from these processes.

[0113] At the same time, this application uses specific terms to describe the embodiments of this application. For example, "one embodiment," "an embodiment," and / or "some embodiments" refer to a certain feature, structure, or characteristic related to at least one embodiment of this application. Therefore, it should be emphasized and noted that "one embodiment," "an embodiment," or "an alternative embodiment" mentioned twice or multiple times in different locations in this specification does not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of this application may be appropriately combined.

[0114] Some aspects of the present application can be performed entirely by hardware, entirely by software (including firmware, resident software, microcode, etc.), or by a combination of hardware and software. The above hardware or software can be referred to as "data blocks", "modules", "engines", "units", "components" or "systems". The processor can be one or more application-specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DAPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), processors, controllers, microcontrollers, microprocessors or combinations thereof. In addition, various aspects of the present application may be expressed as computer products located in one or more computer-readable media, which include computer-readable program code. For example, computer-readable media may include, but are not limited to, magnetic storage devices (e.g., hard disks, floppy disks, tapes...), optical disks (e.g., compact disks CDs, digital versatile disks DVDs...), smart cards, and flash memory devices (e.g., cards, sticks, key drives...).

[0115] A computer-readable medium may include a propagated data signal embodying computer program code, for example, in baseband or as part of a carrier wave. The propagated signal may be in a variety of forms, including electromagnetic, optical, etc., or a suitable combination thereof. A computer-readable medium may be any computer-readable medium other than a computer-readable storage medium that can be connected to an instruction execution system, apparatus, or device to communicate, propagate, or transmit the program for use. The program code on the computer-readable medium may be transmitted via any suitable medium, including radio, cable, fiber optic cable, radio frequency signal, or similar medium, or any combination of the above.

[0116] Similarly, it should be noted that, in order to simplify the presentation of this application and thus facilitate understanding of one or more embodiments of the invention, the foregoing descriptions of the embodiments of this application sometimes combine multiple features into a single embodiment, figure, or description thereof. However, this disclosure method does not mean that the subject matter of this application requires more features than those recited in the claims. In fact, an embodiment may have fewer features than all of the features of a single embodiment disclosed above.

[0117] Although the present application has been described with reference to the current specific embodiments, ordinary technicians in this technical field should recognize that the above embodiments are only used to illustrate the present application, and various equivalent changes or substitutions can be made without departing from the spirit of the present application. Therefore, as long as the changes and modifications to the above embodiments are within the scope of the essential spirit of the present application, they will fall within the scope of the claims of the present application.

Claims

1. A satellite attitude control method for satellite-ground alignment, characterized in that: include: Acquiring output information of a detector of a payload, the output information including coordinates of a light spot of a landmark laser on the detector through an optical lens; Calculating the optical axis deviation angle between the payload reference optical axis and the landmark laser according to the output information; Calculate a final satellite station vector according to the optical axis deflection, where the final satellite station vector is a unit vector pointing from the satellite's center of mass to the ground communication station in the satellite's coordinate system; Calculate and measure the subtending angle according to the final satellite station vector; Obtaining the satellite platform's attitude angle and angular velocity; Calculating a satellite three-axis control torque according to the station angle, the satellite platform attitude angle, and the satellite platform attitude angular velocity, and controlling the satellite platform motion according to the satellite three-axis control torque; The optical axis deviation angle is calculated using the following formula: α=K cam (x-x0) β=K cam (y-y0) Wherein, α and β are the optical axis deviation angles, K cam is a constant coefficient related to the geometric distance and focal length of the payload detector pixel, x0 and y0 are the offset correction values ​​of the payload detector center point, and the coordinates of the light spot are (x, y, 0); Calculating the final star station vector according to the optical axis deflection angle includes: calculating an initial star station vector according to the optical axis deflection angle, the initial star station vector being a vector from the satellite's center of mass to the ground communication station in a detector coordinate system; converting the initial star station vector into an intermediate star station vector according to a transformation matrix from the detector coordinate system to the satellite body coordinate system; and regularizing the intermediate star station vector to obtain the final star station vector. Calculating the satellite three-axis control torque according to the station angle, the satellite platform attitude angle, and the satellite platform attitude angular velocity includes: calculating the station error angle according to the station angle and a preset desired station angle; calculating the station axial error attitude angle according to the satellite platform attitude angle and a preset desired attitude angle; calculating the error angular velocity according to the satellite platform attitude angular velocity and a preset desired angular velocity; and calculating the satellite three-axis control torque according to the station error angle, the station axial error attitude angle, and the error angular velocity. The satellite three-axis control torque is calculated using the following formula: Among them, K p =[k px k py k pz ]、K d =[k dx k dy k dz ]、K i =[k ix k iy k iz ] are the designed proportional, differential and integral control parameters respectively, T i_pre =[T ix_pre T iy_pre T iz_pre ] T T of the previous control cycle i , the initial value is T i0 =[0 0 0] T , station error angle and θ e , the station axial error attitude angle ψ e , error angular velocity ω ex 、ω ey 、ω ez .

2. The method according to claim 1, wherein The initial star station vector is calculated using the following formula: ST m =[-a -b 1] T Wherein, α and β are the optical axis deviation angles.

3. The method according to claim 1, wherein The measured subtended angle is calculated using the following formula: i m =-atan2(μST bx ,μST bz ) Among them, Mesure_A bo is the measured station angle, and the final star station vector is (μST bx , μST by , μST bz ), the function f=atan2(m,n) is used to find the variable The inverse tangent value of , and the range is f∈[-π,π].

4. The method according to claim 1, wherein Also includes: Obtaining a safe operating range of the angular velocity of the star, wherein the safe operating range includes a maximum forward angular velocity and a maximum reverse angular velocity; The station error angle and the station axial error attitude angle are limited according to the safe operating range.

5. A satellite attitude control system, comprising a payload precision tracking system, wherein the satellite attitude control system is used to align the payload reference optical axis with a ground communication station, characterized in that: Also includes: The payload detector is used to collect the coordinates of the light spot on the detector generated by the landmark laser through the optical lens, and send the coordinates of the light spot to the platform control system of the satellite platform; Satellite platforms, including: Attitude sensor, used to measure the attitude angle and angular velocity of the satellite platform; A platform control system, configured to execute the method according to any one of claims 1 to 3.

6. The system according to claim 5, wherein: The satellite platform also includes: a storage unit, which stores a safe operating range of the satellite angular velocity; the platform control system is further used to: obtain the safe operating range, and limit the station error angle and the station axial error attitude angle according to the safe operating range.

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