Spacecraft tracking and photographing processing method based on equatorial mount and related equipment

Abstract

The application provides a spacecraft tracking and shooting processing method based on an equatorial telescope and related equipment. The coordinate value of the spacecraft is calculated every other calculation period when the spacecraft is in a non-blind area, and each calculation period includes W refresh periods. The step value corresponding to each refresh period in the nth calculation period is determined according to the first coordinate value of the spacecraft at T n and the second coordinate value of the spacecraft at T n+1 . In the nth calculation period, the step motor is driven to rotate according to the step value and the refresh period, so that the equatorial telescope tracks and shoots the spacecraft. Thus, the coordinate value of the spacecraft is indirectly obtained at a high frequency and provided to the step motor on the equatorial telescope for driving. The equatorial telescope updates the coordinate value of the field of view at a high frequency, which can avoid the spacecraft from leaving the field of view, the tracking process is smoother, and the video of the spacecraft is more coherent. Therefore, the application improves the video display effect of the spacecraft based on the equatorial telescope tracking and shooting.

Description

Technical Field

[0001] This application relates to the field of astronomical observation technology, and in particular to a spacecraft tracking and imaging processing method and related equipment based on an equatorial mount. Background Technology

[0002] Optical tracking of low- and medium-Earth orbit spacecraft is a crucial task in aerospace telemetry, tracking, and near-Earth satellite communications. Tracking and photographing spacecraft using an equatorial mount requires the mount to acquire the spacecraft's coordinates frequently. However, the computing power module within the equatorial mount often lacks sufficient processing power to acquire the spacecraft's coordinates at such a high frequency. For example, the equatorial mount might acquire the spacecraft's coordinates only every 50ms, then drive its stepper motor to rotate to that coordinate position.

[0003] Because the equatorial mount needs a long interval to acquire the spacecraft's coordinates, the actual situation when tracking and filming a spacecraft is that the spacecraft has already moved to the next trajectory point, while the equatorial mount has not yet started moving to the next trajectory point. Therefore, the tracking and filming footage has a noticeable sense of lag, resulting in poor display quality of spacecraft videos based on equatorial mount tracking and filming. Summary of the Invention

[0004] This application provides a spacecraft tracking and imaging processing method and related equipment based on an equatorial mount, in order to solve the problem of poor display effect of spacecraft videos based on equatorial mount tracking and imaging in the prior art.

[0005] In a first aspect, embodiments of this application provide a spacecraft tracking and imaging processing method based on an equatorial mount, including:

[0006] During the phase when the spacecraft is in the non-blind zone, the coordinate value of the spacecraft is calculated every calculation cycle. Each calculation cycle includes W refresh cycles, where W is an integer greater than 1. The refresh cycle represents the interval between two consecutive drives of the stepper motor in the equatorial mount.

[0007] According to the spacecraft in T n The first coordinate value at time T n+1 The second coordinate value at time T determines the step value corresponding to each refresh cycle within the nth calculation cycle. n The time is the start time of the nth calculation cycle, T. n+1 The time is the start time of the (n+1)th calculation cycle, where n is a positive integer;

[0008] During the nth calculation cycle, the stepper motor is driven to rotate according to the step value and refresh cycle so that the equatorial mount can track and photograph the spacecraft.

[0009] Secondly, embodiments of this application provide a spacecraft tracking and imaging processing method based on an equatorial mount, including:

[0010] During the tracking and imaging process in the nth calculation cycle, a timestamp corresponding to a reference frame, the geographical latitude angle of the equatorial mount, and the spatial position information of the spacecraft are generated and recorded every first cycle.

[0011] The field rotation angle and offset direction of the reference frame are determined based on the geographic latitude angle of the equatorial mount and the spatial position information of the spacecraft.

[0012] Linear fitting is performed based on the field rotation angle and timestamp of two adjacent reference frames to obtain the field rotation angle of the video frame between the two adjacent reference frames;

[0013] Each video frame is rotated according to the field rotation angle of each video frame.

[0014] Thirdly, embodiments of this application provide a spacecraft tracking and imaging processing method based on an equatorial mount, including:

[0015] After the spacecraft enters the blind zone phase, the equatorial mount continues to rotate at its highest hourly angular rate;

[0016] When the equatorial mount continues to rotate at its highest hour angle rate, and the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is less than a first preset value, the equatorial mount begins to reduce its hour angle rate.

[0017] When the equatorial mount reduces its hourly angular rate, if the difference between the hourly angular rate of the equatorial mount and the hourly angular rate of the spacecraft is less than a second preset value, the equatorial mount stops reducing its hourly angular rate.

[0018] When the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is zero, the equatorial mount continues to rotate at the hour angle rate of the spacecraft.

[0019] Fourthly, embodiments of this application provide a spacecraft tracking and imaging processing device based on an equatorial mount, comprising:

[0020] The calculation module calculates the spacecraft's coordinates every calculation cycle when the spacecraft is in a non-blind zone. Each calculation cycle includes W refresh cycles, where W is an integer greater than 1. The refresh cycle represents the interval between two consecutive drives of the stepper motor in the equatorial mount.

[0021] Determine the module, based on the spacecraft in T n The first coordinate value at time T n+1 The second coordinate value at time T determines the step value corresponding to each refresh cycle within the nth calculation cycle. nThe time is the start time of the nth calculation cycle, T. n+1 The time is the start time of the (n+1)th calculation cycle, where n is a positive integer;

[0022] The drive module drives the stepper motor to rotate according to the step value and refresh cycle during the nth calculation cycle, so that the equatorial mount can track and photograph the spacecraft.

[0023] Fifthly, embodiments of this application provide a spacecraft tracking and imaging processing device based on an equatorial mount, comprising:

[0024] During the tracking and imaging process in the nth calculation cycle, the generation module generates and records the timestamp corresponding to a reference frame, the geographical latitude angle of the equatorial mount, and the spatial position information of the spacecraft every first cycle.

[0025] The second determining module determines the field rotation angle and offset direction of the reference frame based on the geographic latitude angle of the equatorial mount and the spatial position information of the spacecraft.

[0026] The acquisition module performs linear fitting based on the field rotation angle and timestamp of two adjacent reference frames to obtain the field rotation angle of the video frame between two adjacent reference frames;

[0027] The rotation module rotates each video frame according to the field rotation angle of each video frame.

[0028] Sixthly, embodiments of this application provide a spacecraft tracking and imaging processing device based on an equatorial mount, comprising:

[0029] The first continuous module, after the spacecraft enters the blind zone phase, the equatorial mount continues to rotate at the highest hourly angular rate;

[0030] The reduction module, when the equatorial mount continues to rotate at its highest hour angle rate, and the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is less than a first preset value, the equatorial mount begins to reduce its hour angle rate;

[0031] The stopping module, when the equatorial mount reduces its hourly angular rate, stops reducing its hourly angular rate when the difference between the hourly angular rate of the equatorial mount and the hourly angular rate of the spacecraft is less than a second preset value;

[0032] The second continuous module, when the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is zero, the equatorial mount continues to rotate at the hour angle rate of the spacecraft.

[0033] In a seventh aspect, embodiments of this application provide a communication device, including: a transceiver, a memory, a processor, and a program stored in the memory and executable on the processor; characterized in that the processor is configured to read the program in the memory to implement the steps of the method as described in any of the preceding claims.

[0034] Eighthly, embodiments of this application provide a readable storage medium for storing a program, characterized in that the program, when executed by a processor, implements the steps of the method as described in any of the preceding claims.

[0035] This application embodiment calculates the spacecraft's coordinates at regular intervals during the non-blind zone phase. Each calculation cycle includes W refresh cycles, where W is an integer greater than 1. The refresh cycle represents the interval between two consecutive drives of the stepper motor within the equatorial mount. Based on the spacecraft's position at T... n The first coordinate value at time T n+1 The second coordinate value at time T determines the step value corresponding to each refresh cycle within the nth calculation cycle. n The time is the start time of the nth calculation cycle, T. n+1 The time is the start time of the (n+1)th calculation cycle, where n is a positive integer. During the nth calculation cycle, the stepper motor is driven to rotate according to the step value and refresh cycle, enabling the equatorial mount to track and photograph the spacecraft. This indirectly achieves high-frequency acquisition of the spacecraft's coordinate values ​​and provides them to the stepper motor on the equatorial mount for driving. The equatorial mount's high-frequency update of its own field-of-view coordinate values ​​reduces the time it takes for the spacecraft to leave the field of view, resulting in more continuous video footage. Therefore, this embodiment improves the display effect of spacecraft video captured by equatorial mount tracking. Attached Figure Description

[0036] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the description of the embodiments of this application will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0037] Figure 1 This is one of the flowcharts illustrating a spacecraft tracking and imaging processing method based on an equatorial mount provided in this application embodiment;

[0038] Figure 2 This is the second flowchart of a spacecraft tracking and imaging processing method based on an equatorial mount provided in this application embodiment;

[0039] Figure 3 This is the third flowchart of a spacecraft tracking and imaging processing method based on an equatorial mount provided in this application embodiment;

[0040] Figure 4 This is one of the structural schematic diagrams of a spacecraft tracking and imaging processing device based on an equatorial mount provided in the embodiments of this application;

[0041] Figure 5 This is a second schematic diagram of a spacecraft tracking and imaging processing device based on an equatorial mount provided in this application embodiment;

[0042] Figure 6 This is the third schematic diagram of a spacecraft tracking and imaging processing device based on an equatorial mount provided in this application embodiment;

[0043] Figure 7 This is a schematic diagram of the structure of a communication device provided in an embodiment of this application. Detailed Implementation

[0044] The technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, not all embodiments. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0045] Unless otherwise defined, the technical or scientific terms used in this application shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The terms "first," "second," and similar terms used in this application do not indicate any order, quantity, or importance, but are merely used to distinguish different components. Terms such as "upper," "lower," "left," and "right" are used only to indicate relative positional relationships; when the absolute position of the described object changes, the relative positional relationship also changes accordingly.

[0046] Reference Figure 1 This application provides a spacecraft tracking and imaging processing method based on an equatorial mount, including:

[0047] Step 101: When the spacecraft is in the non-blind zone, calculate the coordinate value of the spacecraft every calculation cycle. Each calculation cycle includes W refresh cycles, where W is an integer greater than 1. The refresh cycle represents the interval between two consecutive drives of the stepper motor in the equatorial mount.

[0048] It should be understood that a spacecraft being in the non-blind zone phase does not mean that the spacecraft can be observed continuously. The non-blind zone phase means that the spacecraft can be observed by controlling the rotation of the equatorial mount; when the spacecraft is in the blind zone phase, it means that no matter how the rotation of the equatorial mount is controlled, the equatorial mount cannot observe the spacecraft.

[0049] It's important to clarify that calculating the spacecraft's coordinates every calculation cycle doesn't mean a full calculation cycle is required to calculate the spacecraft's coordinates. Rather, the computing module on the equatorial mount needs to calculate the spacecraft's coordinates not only but also schedule and calculate other tasks. Therefore, calculating the spacecraft's coordinates once per calculation cycle means that within the length of one calculation cycle, the computing module on the equatorial mount can only provide time for calculating the spacecraft's coordinates once. For example, if a calculation cycle is 50ms, and the computing module on the equatorial mount takes 5ms to calculate the spacecraft's coordinates, this means that within the 50ms timeframe, the computing module can only provide 5ms to calculate the spacecraft's coordinates; the remaining 45ms is used for scheduling and calculating other tasks.

[0050] It should be noted that each calculation cycle includes W refresh cycles, where W is an integer greater than 1. The refresh cycle represents the interval between two consecutive drives of the stepper motors within the equatorial mount. This means that the time it takes for the equatorial mount to calculate the spacecraft's coordinates is longer than the interval between two consecutive drives of the stepper motors on the equatorial mount. For example, when W is 3, it means there are three refresh cycles within one calculation cycle. In other words, for the equatorial mount to acquire the spacecraft's coordinates once, the stepper motors on the equatorial mount need to receive three step values.

[0051] Step 102, based on the spacecraft at T n The first coordinate value at time T n+1 The second coordinate value at time T determines the step value corresponding to each refresh cycle within the nth calculation cycle. n The time is the start time of the nth calculation cycle, T. n+1 The time is the start time of the (n+1)th calculation cycle, where n is a positive integer;

[0052] It should be noted that the step value is the same for different refresh cycles within a single calculation cycle, but the step value may be different for different refresh cycles within different calculation cycles.

[0053] It should be understood that obtaining information about the spacecraft in T n The first coordinate value at time T n+1 The second coordinate value at a given time can be obtained by using the TLE (Two Lines Elements) data of the spacecraft in orbit and the latitude and longitude of the equatorial mount, by calling the open-source SGDP4 satellite orbit calculation application to obtain the spacecraft's coordinate values ​​at the corresponding time. n The first coordinate value at time T n+1After obtaining the second coordinate value at a given time, the difference between the coordinate values ​​at these two times is obtained. Then, the difference between the coordinate values ​​at these two times is divided by the number of refresh cycles of the stepper motor contained between these two times to obtain the step value corresponding to a single refresh cycle.

[0054] Step 103: Drive the stepper motor to rotate according to the step value and refresh cycle during the nth calculation cycle so that the equatorial mount can track and photograph the spacecraft.

[0055] It should be understood that, as shown in step 102, the step value is the same for all refresh cycles within a calculation cycle. Therefore, in T... n Time to T n+1 Between moments, a fixed step value is provided to the stepper motor on the equatorial mount at the beginning of each refresh cycle.

[0056] Optionally, in some embodiments, the above-described spacecraft tracking and imaging processing method based on an equatorial mount can also be applied to communication between ground antennas and communication satellites.

[0057] In this embodiment, when the spacecraft is in a non-blind zone, the spacecraft's coordinates are calculated every calculation cycle. Each calculation cycle includes W refresh cycles, where W is an integer greater than 1. The refresh cycle represents the interval between two consecutive drives of the stepper motor within the equatorial mount. Based on the spacecraft's position at T... n The first coordinate value at time T n+1 The second coordinate value at time T determines the step value corresponding to each refresh cycle within the nth calculation cycle. n The time is the start time of the nth calculation cycle, T. n+1 The time is the start time of the (n+1)th calculation cycle, where n is a positive integer. During the nth calculation cycle, the stepper motor is driven to rotate according to the step value and refresh cycle, enabling the equatorial mount to track and photograph the spacecraft. This indirectly achieves high-frequency acquisition of the spacecraft's coordinate values ​​and provides them to the stepper motor on the equatorial mount for driving. The equatorial mount's high-frequency update of its own field-of-view coordinate values ​​prevents the spacecraft from leaving the field of view, making the tracking process smoother and more stable, and resulting in more coherent video footage of the spacecraft. Therefore, this embodiment improves the display effect of spacecraft video captured by equatorial mount tracking.

[0058] Optionally, in some embodiments, during the tracking and imaging process in the nth calculation cycle, a timestamp corresponding to a reference frame, the geographic latitude angle of the equatorial mount, and the spacecraft's spatial position information are generated and recorded every first cycle.

[0059] The field rotation angle of the reference frame is determined based on the geographic latitude angle of the equatorial mount and the spatial position information of the spacecraft.

[0060] The field rotation angle of the video frame between two adjacent reference frames is obtained by linear fitting based on the field rotation angle and timestamp of two adjacent reference frames.

[0061] Each video frame is rotated according to the field rotation angle of each video frame.

[0062] It should be noted that this first period is the time interval between two adjacent reference frames, which is also a multiple of the aforementioned calculation period (e.g., 5 times). Further explanation is needed: the reference frame is not a single frame from the actual captured video, but rather a set of additionally recorded reference data. At the beginning of each of the aforementioned first periods, the equatorial mount's calculation module outputs the current timestamp, the equatorial mount's latitude, and the spacecraft coordinates calculated within that calculation period. This data is then used to batch calculate the corresponding field rotation angles, which are then used for subsequent linear interpolation processing of the video frame field rotation angles.

[0063] It should be understood that rotating each video frame according to the field rotation angle of each video frame can be done using MATLAB software by calling the image deflection function imrotate. Based on the video frame number, timestamp, and field rotation angle, each frame image is rotated to restore the upright image in the theodolite mode. The video frame number refers to the order of the video frames generated when the video was captured.

[0064] It should be understood that by using the above settings, each frame of the captured spacecraft video is corrected according to the field rotation angle to obtain the spacecraft video captured in theodolite mode, thereby solving the dynamic field rotation problem that occurs when capturing spacecraft based on the equatorial mount, and further improving the display effect of the spacecraft video captured based on the equatorial mount.

[0065] Optionally, in some embodiments, the field rotation angle of the reference frame is determined using the following formula:

[0066]

[0067] Where η is the field rotation angle, A is the latitude angle of the current location of the equatorial mount in the WGS84 coordinate system, h is the elevation angle in the horizontal coordinate system of the spacecraft's theodolite mode, and A is the altitude angle in the horizontal coordinate system. Z δ is the azimuth angle of the horizontal coordinate system in the spacecraft theodolite mode, and δ is the declination angle of the first equatorial coordinate system in the spacecraft equatorial mode.

[0068] The offset direction is determined based on the spacecraft's position relative to the equatorial mount.

[0069] It should be noted that when the spacecraft is located in the western half of the equatorial mount, the field rotation angle is clockwise, and when the spacecraft is located in the eastern half of the equatorial mount, the field rotation angle is counterclockwise.

[0070] Optionally, in some embodiments, a linear fit is performed based on the field rotation angle and timestamp of two adjacent reference frames to obtain the field rotation angle of the video frame between two adjacent reference frames:

[0071]

[0072] Where T2 and T3 are the shooting times of two adjacent reference frames, and the field rotation angle at time T2 is η. T2 The field rotation angle of the reference frame at time T3 is η. T3 t4 is the time between time T2 and time T3, and η4 is the field rotation angle at time t4.

[0073] Optionally, in some embodiments, the above linear fitting formula can be implemented using the TREND, MATCH, and INDEX fitting functions built into Excel software, ultimately fitting the frame number, timestamp, and field rotation angle data information for each video frame.

[0074] Optionally, in some embodiments, before calculating the coordinate value of the spacecraft every calculation cycle during the phase when the spacecraft is in a non-blind zone, the method further includes: continuously accelerating the motion according to the first time fast forward acceleration during a first preset time period, so that the motion speed of the equatorial mount is greater than the motion speed of the spacecraft.

[0075] During the second preset time period, the motion is continuously decelerated according to the second time fast forward acceleration so that the speed of the equatorial mount is equal to the speed of the spacecraft. The start time of the second preset time period is the end time of the first preset time period, or the start time of the second preset time period is after the end time of the first preset time period.

[0076] It should be noted that in order for the equatorial mount to move faster than the spacecraft, due to the initial time and position difference between the equatorial mount and the spacecraft, the equatorial mount's fastest speed must exceed that of the spacecraft in order to track the spacecraft.

[0077] It should be understood that the absolute values ​​of the first acceleration and the second acceleration can be the same.

[0078] In this embodiment of the application, the above settings enable the equatorial mount to start chasing the spacecraft when its velocity is zero, and the equatorial mount will not overshoot or lose synchronization, thereby improving the tracking accuracy of the equatorial mount and thus improving the display effect of the spacecraft video captured by the equatorial mount.

[0079] Optionally, in some embodiments, based on a first preset time period, a first acceleration, a second preset time period, and a second acceleration, it is determined whether a third preset time period exists. The third preset time period is between the first preset time period and the second preset time period, and the equatorial mount moves at a constant speed during the third preset time period.

[0080] Optionally, in some embodiments, determining whether a third preset time period exists based on a first preset time period, a first acceleration, a second preset time period, and a second acceleration includes:

[0081] When ΔT satisfies In the case of a third preset time period, it is determined that there exists a third preset time period, where ΔT is the time difference between the equatorial mount and the spacecraft when the equatorial mount reaches the previous trajectory point of the spacecraft and the velocity of the equatorial mount is zero, N is the time fast-forward multiple reached by the equatorial mount after the first preset time period, and a is the absolute value of the first acceleration.

[0082] Optionally, in some embodiments, if a third preset time period t2 exists, the hour angle of the equatorial mount is [value missing] within the first preset time period t1. The declination angle of the equatorial mount is During the third preset time period t2, the hour angle of the equatorial mount is... The declination angle of the equatorial mount is During the second preset time period t3, the hour angle of the equatorial mount is The declination angle of the equatorial mount is

[0083] In the absence of t2, the hour angle of the equatorial mount within t1 is... The declination angle of the equatorial mount is Within t3, the hour angle of the equatorial mount is... The declination angle of the equatorial mount is

[0084] Based on the hour angle and declination angle of the equatorial mount, control the rotation of the equatorial mount so that it can track and photograph the spacecraft.

[0085] Optionally, in some embodiments, the hour angle A'(t) and declination angle B'(t) of the equatorial mount are calculated by the SGDP4 satellite orbit calculation application.

[0086] Optionally, in some embodiments, the equatorial mount continues to rotate at its highest hourly angular rate after the spacecraft enters the blind zone phase;

[0087] When the equatorial mount continues to rotate at its highest hour angle rate, and the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is less than a first preset value, the equatorial mount begins to reduce its hour angle rate.

[0088] When the equatorial mount reduces its hourly angular rate, the equatorial mount stops reducing its hourly angular rate when the difference between the equatorial mount's hourly angular rate and the spacecraft's hourly angular rate is less than a second preset value.

[0089] With the equatorial mount ceasing to decrease its hourly angular rate, it continues to rotate at its current hourly angular rate.

[0090] When the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is zero, the equatorial mount continues to rotate at the hour angle rate of the spacecraft.

[0091] In this embodiment of the application, the above settings enable the equatorial mount to quickly pass through the blind zone and track the target without overshooting, thereby improving the display effect of spacecraft video captured by the equatorial mount tracking.

[0092] Optionally, in some embodiments, before the equatorial mount continues to rotate at its highest hourly angular rate after the spacecraft enters the blind zone phase, the following steps are also included:

[0093] Every second period, the equatorial mount calculates the change in the spacecraft's hour angle.

[0094] Calculate the spacecraft’s current angular rate based on the second period and the change in hour angle;

[0095] Based on the spacecraft's current hourly angular rate, determine whether the spacecraft has entered the blind zone phase.

[0096] Reference Figure 2 This application provides a spacecraft tracking and imaging processing method based on an equatorial mount, including:

[0097] Step 201: During the tracking and imaging process in the nth calculation cycle, a timestamp, the geographic latitude angle of the equatorial mount, and the spacecraft's spatial position information are generated and recorded every first cycle for a reference frame.

[0098] Step 202: Determine the field rotation angle and offset direction of the reference frame based on the geographic latitude angle of the equatorial mount and the spacecraft's spatial position information;

[0099] Step 203: Perform linear fitting based on the field rotation angle and timestamp of two adjacent reference frames to obtain the field rotation angle of the video frame between two adjacent reference frames;

[0100] Step 204: Rotate each video frame according to the field rotation angle of each video frame.

[0101] It should be noted that this first period is the time interval between two adjacent reference frames, which is also a multiple of the aforementioned calculation period (e.g., 5 times). Further explanation is needed: the reference frame is not a single frame from the actual captured video, but rather a set of additionally recorded reference data. At the beginning of each first period, the equatorial mount's calculation module outputs the current timestamp, the equatorial mount's latitude, and the spacecraft coordinates calculated for that calculation period. This data is then used to batch calculate the corresponding field rotation angles, which are then used for subsequent linear interpolation processing of the video frame field rotation angles.

[0102] It should be understood that by using the above settings, each frame of the captured spacecraft video is corrected according to the field rotation angle to obtain the spacecraft video captured in theodolite mode, thereby solving the dynamic field rotation problem that occurs when capturing spacecraft based on the equatorial mount, and further improving the display effect of the spacecraft video captured based on the equatorial mount.

[0103] Optionally, in some embodiments, determining the field rotation angle and offset direction of the reference frame based on the geographic latitude angle of the equatorial mount and the observation coordinate information of the spacecraft includes:

[0104] The field spin angle is calculated based on the first formula: First formula:

[0105] Where η is the field rotation angle, A is the latitude angle of the current location of the equatorial mount in the WGS84 coordinate system, h is the elevation angle in the horizontal coordinate system of the spacecraft's theodolite mode, and A is the altitude angle in the horizontal coordinate system. Z δ is the azimuth angle of the horizontal coordinate system in the spacecraft theodolite mode, and δ is the declination angle of the first equatorial coordinate system in the spacecraft equatorial mode.

[0106] The offset direction is determined based on the spacecraft's position information on the equatorial mount.

[0107] Optionally, in some embodiments, obtaining the field rotation angle of a video frame between two adjacent reference frames by performing linear fitting based on the field rotation angle and timestamp of two adjacent reference frames includes:

[0108] The field spin angle is calculated based on the second formula, which is:

[0109]

[0110] Where T2 and T3 are the recording times of two adjacent reference frames, and the field rotation angle at time T2 is η. T2 The field rotation angle of the reference frame at time T3 is η. T3 t4 is the time between time T2 and time T3, η 4 Let t4 be the field rotation angle.

[0111] Reference Figure 3 This application provides a spacecraft tracking and imaging processing method based on an equatorial mount, including:

[0112] Step 301: After the spacecraft enters the blind zone phase, the equatorial mount continues to rotate at its highest hourly angular rate;

[0113] Step 302: When the equatorial mount continues to rotate at its highest hour angular rate, and the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is less than a first preset value, the equatorial mount begins to reduce its hour angular rate.

[0114] Step 303: When the equatorial mount reduces its hourly angular rate, and the difference between the hourly angular rate of the equatorial mount and the hourly angular rate of the spacecraft is less than a second preset value, the equatorial mount stops reducing its hourly angular rate.

[0115] Step 304: When the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is zero, the equatorial mount continues to rotate at the hour angle rate of the spacecraft.

[0116] In this embodiment of the application, the above settings enable the equatorial mount to quickly re-track the spacecraft through the blind zone, and the equatorial mount will not overshoot, thereby improving the display effect of the spacecraft video captured by the equatorial mount.

[0117] Optionally, in some embodiments, before the equatorial mount continues to rotate at its highest hourly angular rate after the spacecraft enters the blind zone phase, the method further includes:

[0118] Every second period, the equatorial mount calculates the change in the spacecraft's hour angle.

[0119] Calculate the spacecraft’s current angular rate based on the second period and the change in hour angle;

[0120] Based on the spacecraft's current hourly angular rate, determine whether the spacecraft has entered the blind zone phase.

[0121] Reference Figure 4 This application provides a spacecraft tracking and imaging processing device based on an equatorial mount, comprising:

[0122] The calculation module 401 calculates the coordinates of the spacecraft every calculation cycle when the spacecraft is in the non-blind zone. Each calculation cycle includes W refresh cycles, where W is an integer greater than 1. The refresh cycle represents the interval between two consecutive drives of the stepper motor in the equatorial mount.

[0123] The first determining module 402, based on the spacecraft at T... n The first coordinate value at time T n+1 The second coordinate value at time T determines the step value corresponding to each refresh cycle within the nth calculation cycle. nThe time is the start time of the nth calculation cycle, T. n+1 The time is the start time of the (n+1)th calculation cycle, where n is a positive integer;

[0124] The drive module 403 drives the stepper motor to rotate according to the step value and refresh cycle in the nth calculation cycle so that the equatorial mount can track and photograph the spacecraft.

[0125] Reference Figure 5 This application provides a spacecraft tracking and imaging processing device based on an equatorial mount, comprising:

[0126] During the tracking and imaging process in the nth calculation cycle, the generation module 501 generates and records the timestamp, the geographical latitude angle of the equatorial mount, and the spacecraft's spatial position information corresponding to a reference frame every first cycle.

[0127] The second determining module 502 determines the field rotation angle and offset direction of the reference frame based on the geographic latitude angle of the equatorial mount and the observation coordinate information of the spacecraft;

[0128] The module 503 obtains the field rotation angle of the video frame between two adjacent reference frames by performing linear fitting based on the field rotation angle and timestamp of two adjacent reference frames.

[0129] The rotation module 504 rotates each video frame according to the field rotation angle of each video frame.

[0130] Reference Figure 6 This application provides a spacecraft tracking and imaging processing device based on an equatorial mount, comprising:

[0131] The first continuous module 601 rotates the equatorial mount at the highest hourly angular rate after the spacecraft enters the blind zone phase.

[0132] The lowering module 602, when the equatorial mount continues to rotate at its highest hour angle rate, when the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is less than a first preset value, the equatorial mount begins to reduce its hour angle rate.

[0133] Stop module 603, when the equatorial mount reduces its hourly angular rate, when the difference between the hourly angular rate of the equatorial mount and the hourly angular rate of the spacecraft is less than a second preset value, the equatorial mount stops reducing its hourly angular rate.

[0134] The second continuous module 604 rotates continuously at the time angular rate of the spacecraft when the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is zero.

[0135] Reference Figure 7This application also provides a communication device. The communication device may include a processor 701, a memory 702, and a program 7021 stored in the memory 702 and executable on the processor 701.

[0136] When program 7021 is executed by processor 701, it can achieve the following: Figure 1 Any steps in the corresponding method embodiments and the achievement of the same beneficial effects will not be repeated here.

[0137] Those skilled in the art will understand that all or part of the steps of the methods described in the above embodiments can be implemented by hardware related to program instructions, and the program can be stored in a readable medium.

[0138] This application also provides a readable storage medium storing a computer program, which, when executed by a processor, can perform the above-described functions. Figure 1 Any step in the corresponding method embodiment can achieve the same technical effect, and will not be repeated here to avoid repetition.

[0139] The computer-readable storage medium of this application embodiment can be any combination of one or more computer-readable media. The computer-readable medium can be a computer-readable signal medium or a computer-readable storage medium. For example, a computer-readable storage medium can be an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination thereof. More specific examples of computer-readable storage media (a non-exhaustive list) include: an electrical connection having one or more wires, a portable computer disk, a hard disk, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage device, magnetic storage device, or any suitable combination thereof. In this document, a computer-readable storage medium can be any tangible medium containing or storing a program that can be used by or in conjunction with an instruction execution system, apparatus, or device.

[0140] Computer-readable signal media may include data signals propagated in baseband or as part of a carrier wave, carrying computer-readable program code. Such propagated data signals may take various forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination thereof. Computer-readable signal media may also be any computer-readable medium other than computer-readable storage media, capable of sending, propagating, or transmitting programs for use by or in connection with an instruction execution system, apparatus, or device.

[0141] The program code contained on the storage medium can be transmitted using any suitable medium, including but not limited to wireless, wire, optical fiber, RF, etc., or any suitable combination thereof.

[0142] Computer program code for performing the operations of this application can be written in one or more programming languages ​​or a combination thereof. Programming languages ​​include object-oriented programming languages—such as Java, Smalltalk, and C++—as well as conventional procedural programming languages—such as the "C" language or similar programming languages. The program code can be executed entirely on the user's computer, partially on the user's computer, as a standalone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or terminal. In cases involving remote computers, the remote computer can be connected to the user's computer via any type of network—including a local area network (LAN) or a wide area network (WAN)—or can be connected to an external computer (e.g., via the Internet using an Internet service provider).

[0143] The above are preferred embodiments of the present application. It should be noted that, for those skilled in the art, several improvements and modifications can be made without departing from the principles of the present application, and these improvements and modifications should also be considered within the scope of protection of the present application.

Claims

1. A spacecraft tracking and imaging processing method based on an equatorial mount, characterized in that, include: During the phase when the spacecraft is in a non-blind zone, the coordinate value of the spacecraft is calculated every calculation cycle. Each calculation cycle includes W refresh cycles, where W is an integer greater than 1. The refresh cycle represents the interval between two consecutive drives of the stepper motor in the equatorial mount. According to the spacecraft in T n The first coordinate value at time T n+1 The second coordinate value at time T determines the step value corresponding to each refresh cycle within the nth calculation cycle. n The time is the start time of the nth calculation cycle, and T is... n+1 The time is the start time of the (n+1)th calculation cycle, where n is a positive integer; During the nth calculation cycle, the stepper motor is driven to rotate according to the step value and the refresh cycle, so that the equatorial mount tracks and photographs the spacecraft; During the tracking and imaging process in the nth calculation cycle, a timestamp corresponding to a reference frame, the geographic latitude angle of the equatorial mount, and the observation coordinate information of the spacecraft are generated and recorded every first cycle. The field rotation angle and offset direction of the reference frame are determined based on the geographic latitude angle of the equatorial mount and the observation coordinate information of the spacecraft. Linear fitting is performed based on the field rotation angle and timestamp of two adjacent reference frames to obtain the field rotation angle of the video frame between the two adjacent reference frames; Rotate each video frame according to the field rotation angle of each video frame.

2. The spacecraft tracking and imaging processing method based on an equatorial mount according to claim 1, characterized in that, The determination of the field rotation angle and offset direction of the reference frame based on the geographic latitude angle of the equatorial mount and the observation coordinate information of the spacecraft includes: The field rotation angle is calculated based on the first formula, where: ,in, For the field rotation angle, The latitude and longitude angle of the equatorial mount in the WGS84 coordinate system is the current location. The elevation angle in the horizon coordinate system under the theodolite mode of the spacecraft. The azimuth angle in the horizon coordinate system of the spacecraft in theodolite mode. The declination angle of the first equatorial coordinate system in the equatorial mode of the spacecraft; The offset direction is determined based on the spacecraft's position information on the equatorial mount.

3. The spacecraft tracking and imaging processing method based on an equatorial mount according to claim 1, characterized in that, The step of obtaining the field rotation angle of the video frame between two adjacent reference frames by performing linear fitting based on the field rotation angle and timestamp of two adjacent reference frames includes: The field rotation angle is calculated based on the second formula, which is: ; Where T2 and T3 are the recording times of two adjacent reference frames, and the field rotation angle at time T2 is... The field rotation angle of the reference frame at time T3 is t4 is the time between time T2 and time T3. Let t4 be the field rotation angle.

4. The spacecraft tracking and imaging processing method based on an equatorial mount according to claim 1, characterized in that, Before calculating the spacecraft's coordinates every calculation cycle during the phase when the spacecraft is in a non-blind zone, the method further includes: During a first preset time period, the motion is continuously accelerated according to a first time fast forward acceleration so that the motion speed of the equatorial mount is greater than the motion speed of the spacecraft. During a second preset time period, the motion is continuously decelerated according to a second time fast-forward acceleration so that the motion speed of the equatorial mount is equal to the motion speed of the spacecraft. The start time of the second preset time period is the end time of the first preset time period, or the start time of the second preset time period is after the end time of the first preset time period.

5. The spacecraft tracking and imaging processing method based on an equatorial mount according to claim 4, characterized in that, include: Based on a first preset time period, a first time-forward acceleration, a second preset time period, and a second time-forward acceleration, it is determined whether a third preset time period exists. The third preset time period is between the first preset time period and the second preset time period. During the third preset time period, the equatorial mount moves at a constant speed according to a preset time-forward multiple. The preset time-forward multiple is determined based on the first preset time period and the first time-forward multiple.

6. The spacecraft tracking and imaging processing method based on an equatorial mount according to claim 5, characterized in that, The determination of whether a third preset time period exists based on a first preset time period, a first time-lapse acceleration, a second preset time period, and a second time-lapse acceleration includes: exist satisfy In the case of the aforementioned third preset time period, it is determined that such a time period exists. The time difference between the equatorial mount and the spacecraft when the equatorial mount reaches the previous trajectory point of the spacecraft and the velocity of the equatorial mount is zero, where N is the time fast-forward multiple achieved by the equatorial mount after the first preset time period, and a is the absolute value of the first time fast-forward acceleration.

7. The spacecraft tracking and imaging processing method based on an equatorial mount according to claim 1, characterized in that, include: After the spacecraft enters the blind zone phase, the equatorial mount continues to rotate at its highest hourly angular rate; When the equatorial mount continues to rotate at its highest hour angle rate, and the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is less than a first preset value, the equatorial mount begins to reduce its hour angle rate. When the equatorial mount reduces its hourly angular rate, if the difference between the hourly angular rate of the equatorial mount and the hourly angular rate of the spacecraft is less than a second preset value, the equatorial mount stops reducing its hourly angular rate. When the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is zero, the equatorial mount continues to rotate at the hour angle rate of the spacecraft.

8. The spacecraft tracking and imaging processing method based on an equatorial mount according to claim 7, characterized in that, The procedure further includes the following steps: after the spacecraft enters the blind zone phase, and before the equatorial mount continues to rotate at its highest hourly angular rate: Every second cycle, the equatorial mount calculates the change in the spacecraft's hour angle. Based on the second period and the change in time angle, the current time angle rate of the spacecraft is calculated; Based on the spacecraft's current hourly angular rate, determine whether the spacecraft has entered a blind zone phase.

9. A spacecraft tracking and imaging processing device based on an equatorial mount, characterized in that, include: The calculation module calculates the coordinates of the spacecraft every calculation cycle when the spacecraft is in a non-blind zone. Each calculation cycle includes W refresh cycles, where W is an integer greater than 1. The refresh cycle represents the interval between two consecutive drives of the stepper motor in the equatorial mount. The first determining module, based on the spacecraft's position at T... n The first coordinate value at time T n+1 The second coordinate value at time T determines the step value corresponding to each refresh cycle within the nth calculation cycle. n The time is the start time of the nth calculation cycle, and T is... n+1 The time is the start time of the (n+1)th calculation cycle, where n is a positive integer; The drive module drives the stepper motor to rotate according to the step value and the refresh cycle during the nth calculation cycle, so that the equatorial mount can track and photograph the spacecraft. During the tracking and imaging process in the nth calculation cycle, the generation module generates and records the timestamp, the geographical latitude angle of the equatorial mount, and the spacecraft's spatial position information for a reference frame every first cycle. The second determining module determines the field rotation angle and offset direction of the reference frame based on the geographic latitude angle of the equatorial mount and the observation coordinate information of the spacecraft; The acquisition module performs linear fitting based on the field rotation angle and timestamp of two adjacent reference frames to obtain the field rotation angle of the video frame between two adjacent reference frames; The rotation module rotates each video frame according to the field rotation angle of each video frame.

10. The apparatus according to claim 9, characterized in that, The device further includes: The first continuous module, after the spacecraft enters the blind zone phase, the equatorial mount continues to rotate at the highest hourly angular rate; The reduction module, when the equatorial mount continues to rotate at its highest hour angle rate, and the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is less than a first preset value, the equatorial mount begins to reduce its hour angle rate; The stopping module, when the equatorial mount reduces its hourly angular rate, stops reducing its hourly angular rate when the difference between the hourly angular rate of the equatorial mount and the hourly angular rate of the spacecraft is less than a second preset value; The second continuous module, when the difference between the hour angle of the equatorial mount and the hour angle of the spacecraft is zero, the equatorial mount continues to rotate at the hour angle rate of the spacecraft.

11. A communication device, comprising: A transceiver, a memory, a processor, and a program stored in the memory and executable on the processor; characterized in that the processor is configured to read the program from the memory to implement the steps of the method as described in any one of claims 1 to 8.

12. A readable storage medium for storing a program, characterized in that, When the program is executed by the processor, it implements the steps of the method as described in any one of claims 1 to 8.

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