Spacecraft orbit transfer phase attitude conversion time sequence generation method and system

Abstract

The application provides a spacecraft orbit transfer section attitude conversion time sequence generation method and system, including the following steps: step S1: obtaining a satellite multiple orbit transfer time sequence; step S2: determining a satellite attitude conversion time during an orbit transfer section according to the satellite multiple orbit transfer time sequence; step S3: compiling an attitude conversion semantic matrix according to the attitude conversion time; step S4: converting the attitude conversion semantic matrix into a real attitude time sequence matrix; and step S5: checking the real attitude time sequence matrix. The application uses a simple matrix to represent an attitude time section, and further converts the attitude time section into a real attitude time section through logical judgment. The generation of the attitude time sequence can be quickly realized by computer programming, and the application has strong practicability in launch window calculation and the like.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of the generation process of attitude time sequence in the satellite orbit transfer segment, in particular to a spacecraft orbit transfer segment attitude conversion time sequence generation method and system. BACKGROUND

[0002] For a spacecraft that needs to change orbit, the attitude mode before and after orbit transfer and during orbit transfer is not the same. Generally, the attitude of the orbit transfer segment is described as a natural language description relative to the firing point, for example, from 110 minutes before the first firing to 40 minutes before the first firing, it is a ground pointing attitude. How to convert this semantic description into a real attitude time sequence, that is, the time from the beginning to the end of the natural time description of the ground attitude, is a problem.

[0003] Patent document one: Chinese invention patent document CN102176163A discloses a method for determining the duration of a task observation, which introduces a method for judging the observable time by calculating the attitude angle pointing to the target. This method is used for task planning and scheduling by calculating the observable time of the ground target.

[0004] Patent document two: Chinese invention patent document CN102073495A discloses a time sequence control method for integrated star computer and attitude, which introduces a method for attitude time sequence control in star software. This method collects attitude data according to a certain time step, calculates attitude parameters and outputs.

[0005] Patent document three: Chinese invention patent document CN106125745A discloses a satellite attitude control method for tracking imaging of a space cooperative target, which introduces an attitude control method for tracking a cooperative target. This method calculates the expected attitude according to the given orbit information, and controls the flywheel to generate control torque for feedback control by the controller.

[0006] Patent document four: Chinese invention patent document CN110827315A discloses a target spacecraft trajectory identification method based on time sequence information, which includes: step 1, window processing of the sequence images formed by the observation sensor of the target spacecraft through prior information; step 2, selecting a marker star to obtain the motion characteristics common to the star; step 3, moving the sequence images as a whole, and projecting the maximum value of the sequence images after moving the whole; step 4, obtaining the trajectory information of the target spacecraft based on the time sequence information; step 5, obtaining the position of the target in the sequence images, and further predicting the approximate position of the target in the next frame to assist the target detection in the next frame image.

[0007] In view of the above related art, the inventors believe that there is a need for multiple orbit transfer of a spacecraft, and the attitude needs to be converted multiple times before and after the orbit transfer. In this process, the semantic description of the attitude time period cannot be simply converted into the real attitude. SUMMARY

[0008] In view of the defects in the prior art, the purpose of the present application is to provide a spacecraft orbit transfer segment attitude conversion time sequence generation method and system.

[0009] According to the spacecraft orbit transfer segment attitude conversion time sequence generation method provided by the present application, the following steps are included:

[0010] Step S1: Obtain the time sequence of multiple orbit transfers of a satellite;

[0011] Step S2: Determine the attitude conversion time of the satellite during the orbit transfer segment according to the time sequence of the multiple orbit transfers of the satellite;

[0012] Step S3: Write an attitude conversion semantic matrix according to the attitude conversion time;

[0013] Step S4: Convert the attitude conversion semantic matrix into a real attitude time sequence matrix;

[0014] Step S5: Check the real attitude time sequence matrix.

[0015] Preferably, in the step S1, the time of multiple firings is obtained from the orbit transfer strategy of a medium-high orbit satellite, and the time of firing includes the start time and the end time.

[0016] Preferably, in the step S2, the conversion time of the satellite attitude relative to the firing segment is determined, the attitude before the Nth firing is determined, and the attitude after the Nth firing is determined.

[0017] The firing segment is the time period from the start of the firing to the end of the firing.

[0018] Preferably, in the step S3, the attitude conversion time sequence relative to the firing segment is used to describe the time period before and after the relative firing segment, and then the attitude conversion semantic matrix is written.

[0019] According to the different attitude conversion times, there are different semantic matrix writing methods. The form of the semantic matrix is [n1, n2, t1, t2], where n1 and n2 represent the number of firings, and t1 and t2 represent the time period relative to the firing time.

[0020] When n1 is not nan and n2 is not nan, t1 is the time relative to the n1th ignition, and t2 is the time relative to the n2th ignition, when t1 and t2 are negative, it means before ignition, when t1 and t2 are positive, it means after ignition; wherein, nan represents that the numerical value does not exist, and does not participate in operation;

[0021] When n1 is not nan and n2 is nan, t1 and t2 are both the time relative to the n1th ignition, when t1 and t2 are negative, it means before ignition, when t1 and t2 are positive, it means after ignition;

[0022] When n1 is nan and n2 is not nan, t1 and t2 are both the time relative to the n2th ignition, when t1 and t2 are negative, it means before ignition, when t1 and t2 are positive, it means after ignition.

[0023] Preferably, in the step S4, the posture conversion semantic matrix is converted into a real posture time sequence, and the posture time sequence is calculated according to a plurality of different semantic matrix types.

[0024] Preferably, in the step S5, the start of one posture time sequence is compared with the end of the previous posture time sequence, if consistent, it is correct, if inconsistent, it is incorrect.

[0025] According to the spacecraft orbit transfer segment posture conversion time sequence generation system provided by the application, the following modules are included:

[0026] Module M1: obtaining the time sequence of multiple orbit transfers of a satellite;

[0027] Module M2: determining the posture conversion time of the satellite during the orbit transfer segment according to the time sequence of multiple orbit transfers of the satellite;

[0028] Module M3: writing a posture conversion semantic matrix according to the posture conversion time;

[0029] Module M4: converting the posture conversion semantic matrix into a real posture time sequence matrix;

[0030] Module M5: checking the real posture time sequence matrix.

[0031] Preferably, in the module M1, the time of multiple ignitions is obtained from the orbit transfer strategy of a medium-high orbit satellite, and the time of ignition includes the start time and the end time.

[0032] Preferably, in the module M2, the posture conversion time of the satellite relative to the ignition segment is determined, the posture before the Nth ignition is determined, and the posture after the Nth ignition is determined;

[0033] The ignition segment is a time period from the start of ignition to the end of ignition.

[0034] Preferably, in the module M3, according to the attitude conversion timing of the relative ignition phase, the attitude conversion timing is described using the time period before and after the relative ignition phase, and then the attitude conversion semantic matrix is written;

[0035] According to the attitude conversion timing, there are different semantic matrix writing methods, and the form of the semantic matrix is [n1, n2, t1, t2], wherein n1 and n2 represent the number of ignition, t1 and t2 represent the time period of the relative ignition timing;

[0036] When n1 is not nan and n2 is not nan, t1 is the timing of the relative n1th ignition, and t2 is the timing of the relative n2th ignition, when t1 and t2 are negative, it means before ignition, and when t1 and t2 are positive, it means after ignition; wherein, nan represents that the numerical value does not exist, and does not participate in operation;

[0037] When n1 is not nan and n2 is nan, t1 and t2 are both the timing of the relative n1th ignition, when t1 and t2 are negative, it means before ignition, and when t1 and t2 are positive, it means after ignition;

[0038] When n1 is nan and n2 is not nan, t1 and t2 are both the timing of the relative n2th ignition, when t1 and t2 are negative, it means before ignition, and when t1 and t2 are positive, it means after ignition.

[0039] Compared with the prior art, the present application has the following beneficial effects:

[0040] 1. Based on the demand for multiple orbit changes of the spacecraft, and the actual situation that the attitude needs to be converted multiple times before and after the orbit change, the semantic description of the attitude time period cannot be simply converted into the real attitude in this process, the attitude time period is represented by a simple matrix, and further converted into the real attitude time period through logical judgment, the generation of the attitude timing can be realized quickly by computer programming, and the method has strong practicability in launch window calculation and the like;

[0041] 2. The attitude conversion semantic matrix used in the present application can be effectively used in programming languages, and can be flexibly modified, and different satellite attitude conversion timings can be more conveniently modified;

[0042] 3. The method provided by the present application can quickly obtain the satellite full-process attitude conversion timing, which can be used for other purposes, for example, the attitude conversion timing is input into the STK software in the way of joint simulation of Matlab software and STK software, and the conversion of the satellite attitude timing is simulated. BRIEF DESCRIPTION OF DRAWINGS

[0043] Other features, objects, and advantages of the application will become more apparent from the following detailed description of non-limiting embodiments thereof, when read in conjunction with the accompanying drawings:

[0044] Figure 1 A flowchart of a method for generating an attitude conversion time sequence of a spacecraft during a variable orbit segment. DETAILED DESCRIPTION

[0045] The application will be described in detail below with specific embodiments. The following examples will help those skilled in the art to further understand the application, but do not limit the application in any form. It should be noted that for those skilled in the art, without departing from the concept of the application, a number of changes and improvements can be made. These are within the scope of the application.

[0046] The embodiment of the application discloses a method for generating an attitude conversion time sequence of a spacecraft during a variable orbit segment, as shown in the embodiment of the application, the flowchart comprises: Figure 1

[0047] Step S1: Obtain the time sequence of multiple variable orbits of the satellite, i.e. the start and end time of the satellite engine ignition. Referring to Table 1, the accurate time of multiple ignitions is obtained from the variable orbit strategy of the medium-high orbit satellite, and the time of ignition includes the start time and the end time.

[0048] Table 1: Example of variable orbit time period

[0049] Start, seconds End, seconds 51623.138 54061.511 128612.229 130971.556 237627.528 238432.345

[0050] The time in the above table is represented as the satellite flight time second count from the satellite-rocket separation time.

[0051] Step S2: Determine the attitude conversion time of the satellite during the variable orbit segment according to the time sequence of multiple variable orbits of the satellite. The attitude conversion time of the satellite during the variable orbit segment is determined, and the attitude conversion time is the time when the satellite converts from one attitude mode to another attitude mode. The time is artificially specified, and is generally described using the relative ignition time, for example, the attitude conversion is performed 140 minutes before the ignition starts. The satellite attitude conversion time relative to the ignition segment is determined, and the attitude before the Nth ignition and the attitude after the Nth ignition are determined; the ignition segment is the time period from the ignition start to the end. That is, what attitude should be before the Nth ignition, and what attitude should be after the Nth ignition.

[0052] Step S3: Refer to Table 2, and write the attitude conversion semantic matrix according to the attitude conversion time. According to the attitude conversion time sequence relative to the ignition segment, the ignition segment is the time period from the ignition start to the end, and the attitude conversion time sequence is described using the time period before and after the relative ignition segment, for example, the cruise attitude is converted to the earth-pointing attitude 140 minutes before the ignition starts, and the earth-pointing attitude is converted to the cruise attitude 70 minutes after the ignition ends.​

[0053] Next, the attitude transition semantic matrix is ​​compiled. Depending on the attitude transition moment, different semantic matrices can be compiled. The general form of the semantic matrix is ​​[n1, n2, t1, t2], where n1 and n2 represent the ignition number, and t1 and t2 represent the time period relative to the ignition moment (ignition time). For example, t1 = -140 and t2 = 70 means the time period is from 140 minutes before the n1 ignition to 70 minutes after the n2 ignition.

[0054] When n1 is not nan (indicating that the value does not exist and is not involved in the calculation) and n2 is not nan, t1 is the time relative to the n1th ignition, and t2 is the time relative to the n2th ignition. When t1 and t2 are negative, it indicates before ignition, and when t1 and t2 are positive, it indicates after ignition. Among them, nan indicates that the value does not exist and is not involved in the calculation.

[0055] When n1 is not nan and n2 is nan, t1 and t2 are both relative to the time of the n1th ignition. When t1 and t2 are negative, it indicates before ignition, and when t1 and t2 are positive, it indicates after ignition.

[0056] When n1 is nan and n2 is not nan, t1 and t2 are both relative to the time of the n2th ignition. When t1 and t2 are negative, it indicates before ignition, and when t1 and t2 are positive, it indicates after ignition.

[0057] Table 2 Example of posture semantic description matrix

[0058] n1 n2 t1 t2 Pose 0 1 0 -140 1 1 2 160 -90 1 2 3 160 -100 1 nan 3 160 inf 1 1 nan -140 -30 2 nan 1 0 160 2 2 nan -90 -30 2 nan 2 0 160 2 3 nan -100 -30 2 nan 3 0 160 2 1 1 -30 0 3 2 2 -30 0 4 3 3 -30 0 5

[0059] In the above table, inf stands for infinity, which means that the satellite is ignited only three times, and the satellite's attitude continues to any time after the third ignition.

[0060] Step S4: Referring to Table 3, convert the posture conversion semantic matrix into a real posture timing matrix. Convert the posture conversion semantic matrix into a real posture timing matrix, and calculate the posture timing according to the three different semantic matrix types. Convert the posture semantic description matrix into a real posture timing matrix.

[0061] Table 3. Example of real posture timing

[0062] Start End Pose 0 43223.14 1 43223.14 49823.14 2 49823.14 54061.51 3 54061.51 63661.51 2 63661.51 123212.2 1 123212.2 126812.2 2 126812.2 130971.6 4 130971.6 140571.6 2 140571.6 231627.5 1 231627.5 235827.5 2 235827.5 238432.3 5 238432.3 248032.3 2 248032.3 Inf 1

[0063] Step S5: Check the correctness of the real posture timing matrix. Compare the start of a posture timing with the end of the previous posture timing. If they are consistent, it is correct; if not, it is incorrect.

[0064] A spacecraft trajectory change phase attitude transition time series generation system provided by the present invention includes the following modules:

[0065] Module M1: Obtain the time sequence of multiple orbit changes of the satellite. Obtain the time of multiple firings from the orbit change strategy of the medium-high orbit satellite, and the time of the firing includes the start time and the end time.

[0066] Module M2: Determine the attitude conversion time of the satellite during the orbit change period according to the time sequence of multiple orbit changes of the satellite. Determine the attitude conversion time of the satellite relative to the firing period, determine the attitude before the Nth firing, and determine the attitude after the Nth firing; the firing period is the time period from the start of the firing to the end of the firing.

[0067] Module M3: Write the attitude conversion semantic matrix according to the attitude conversion time. Write the attitude conversion semantic matrix according to the attitude conversion time sequence relative to the firing period, and the attitude conversion time sequence is described using the time period before and after the relative firing period, and then write the attitude conversion semantic matrix.

[0068] According to the different attitude conversion time, there are different semantic matrix writing methods, and the form of the semantic matrix is [n1, n2, t1, t2], wherein n1 and n2 represent the number of firings, and t1 and t2 represent the time period relative to the firing time.

[0069] When n1 is not nan and n2 is not nan, t1 is the time relative to the n1th firing, and t2 is the time relative to the n2th firing, when t1 and t2 are negative, it means before the firing, and when t1 and t2 are positive, it means after the firing; wherein, nan represents that the numerical value does not exist, and does not participate in the operation.

[0070] When n1 is not nan and n2 is nan, t1 and t2 are both the time relative to the n1th firing, when t1 and t2 are negative, it means before the firing, and when t1 and t2 are positive, it means after the firing.

[0071] When n1 is nan and n2 is not nan, t1 and t2 are both the time relative to the n2th firing, when t1 and t2 are negative, it means before the firing, and when t1 and t2 are positive, it means after the firing.

[0072] Module M4: Convert the attitude conversion semantic matrix into a real attitude time sequence matrix. Convert the attitude conversion semantic matrix into a real attitude time sequence, and calculate the attitude time sequence according to a plurality of different semantic matrix types.

[0073] Module M5: Check the real attitude time sequence matrix. Compare the start of an attitude time sequence with the end of the previous attitude time sequence, if consistent, it is correct, if inconsistent, it is incorrect.

[0074] The difference between the present application and the patent document one is that the attitude is maneuvered in a predetermined manner, and the predetermined attitude period is converted from a natural language description to a natural time description.

[0075] The present application differs from the second patent document in that it does not participate in the posture data acquisition and control parameter calculation, but only uses the predetermined posture conversion timing for solving.

[0076] The present application differs from the third patent document in that it does not participate in the torque controller, and does not involve posture control, but is only used for the calculation of posture timing.

[0077] The present application uses a posture period semantic description matrix to represent the natural language description of the posture period, and further converts it into a natural time description of the posture period through logical judgment of the program.

[0078] Those skilled in the art know that, in addition to implementing the system and each device, module and unit thereof provided by the present application in the form of pure computer readable program code, the same function can also be achieved by logically programming the method steps in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers and embedded microcontrollers. Therefore, the system and each device, module and unit thereof provided by the present application can be considered as a hardware component, and the devices, modules and units included therein for achieving various functions can also be considered as structures within the hardware component; the devices, modules and units for achieving various functions can also be considered as both software modules for achieving methods and structures within the hardware component.

[0079] The specific embodiments of the present application are described above. It should be understood that the present application is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essential content of the present application. The embodiments of the present application and the features in the embodiments can be arbitrarily combined with each other without conflict.

Claims

1. A method for generating a time series of attitude transitions during a spacecraft orbit change phase, characterized in that: The steps include: Step S1: Obtain the time series of multiple satellite orbit changes; Step S2: determining the attitude conversion moment of the satellite during the orbit change period according to the time sequence of the satellite's multiple orbit changes; Step S3: compile a posture transition semantic matrix according to the posture transition moment; Step S4: converting the posture conversion semantic matrix into the real posture timing matrix; Step S5: Check the real posture timing matrix.

2. The method for generating a time series of attitude transition during a spacecraft orbit change according to claim 1, wherein: In step S1, the times of multiple ignitions are obtained from the orbit change strategy of the medium and high orbit satellite, and the ignition times include a start time and an end time.

3. The method for generating a time series of attitude transition during a spacecraft orbit change phase according to claim 1, wherein: In step S2, the transition moment of the satellite attitude relative to the ignition phase is determined, and the attitude before the Nth ignition and the attitude after the Nth ignition are determined; The ignition period is the time period from the start to the end of ignition.

4. The method for generating a time series of attitude transition during a spacecraft orbit change phase according to claim 1, wherein: In the step S3, according to the posture conversion timing relative to the ignition segment, the posture conversion timing is described using the time period before and after the relative ignition segment, and then the posture conversion semantic matrix is ​​compiled; There are different ways to write semantic matrices according to the different posture transition moments. The form of the semantic matrix is ​​[n1, n2, t1, t2], where n1 and n2 represent the ignition times, and t1 and t2 represent the time period relative to the ignition moment. When n1 is not nan and n2 is not nan, t1 is the time relative to the n1th ignition, and t2 is the time relative to the n2th ignition. When t1 and t2 are negative, it indicates before ignition, and when t1 and t2 are positive, it indicates after ignition. Among them, nan indicates that the value does not exist and is not involved in the calculation; When n1 is not nan and n2 is nan, t1 and t2 are both relative to the time of the n1th ignition. When t1 and t2 are negative, it means before ignition, and when t1 and t2 are positive, it means after ignition; When n1 is nan and n2 is not nan, t1 and t2 are both relative to the time of the n2th ignition. When t1 and t2 are negative, it indicates before ignition, and when t1 and t2 are positive, it indicates after ignition.

5. The method for generating a time series of attitude transition during a spacecraft orbit change phase according to claim 1, wherein: In the step S4: the posture conversion semantic matrix is ​​converted into a real posture time sequence, and the posture time sequence is calculated according to a plurality of different semantic matrix types.

6. The method for generating a time series of attitude transition during a spacecraft orbit change phase according to claim 1, wherein: In step S5: the start of a gesture sequence is compared with the end of the previous gesture sequence. If they are consistent, it is correct; if they are inconsistent, it is incorrect.

7. A spacecraft trajectory change phase attitude conversion time series generation system, characterized by: Includes the following modules: Module M1: Obtain the time series of multiple satellite orbit changes; Module M2: Determine the attitude transition moment of the satellite during the orbit change period based on the time sequence of multiple orbit changes of the satellite; Module M3: Write the posture transition semantic matrix according to the posture transition moment; Module M4: Convert the posture transition semantic matrix into the real posture timing matrix; Module M5: Check the real posture timing matrix.

8. The spacecraft trajectory change phase attitude conversion time series generation system according to claim 7, characterized in that: In the module M1, the times of multiple ignitions are obtained from the orbit change strategy of the medium and high orbit satellite, and the ignition times include the start time and the end time.

9. The spacecraft trajectory change phase attitude conversion time series generation system according to claim 7, characterized in that: In the module M2, the transition moment of the satellite attitude relative to the ignition phase is determined, and the attitude before the Nth ignition and the attitude after the Nth ignition are determined; The ignition period is the time period from the start to the end of ignition.

10. The spacecraft trajectory change phase attitude conversion time series generation system according to claim 7, characterized in that: In the module M3, according to the posture conversion timing relative to the ignition segment, the posture conversion timing is described using the time period before and after the relative ignition segment, and then the posture conversion semantic matrix is ​​compiled; There are different ways to write semantic matrices according to the different posture transition moments. The form of the semantic matrix is ​​[n1, n2, t1, t2], where n1 and n2 represent the ignition times, and t1 and t2 represent the time period relative to the ignition moment. When n1 is not nan and n2 is not nan, t1 is the time relative to the n1th ignition, and t2 is the time relative to the n2th ignition. When t1 and t2 are negative, it indicates before ignition, and when t1 and t2 are positive, it indicates after ignition. Among them, nan indicates that the value does not exist and is not involved in the calculation; When n1 is not nan and n2 is nan, t1 and t2 are both relative to the time of the n1th ignition. When t1 and t2 are negative, it means before ignition, and when t1 and t2 are positive, it means after ignition; When n1 is nan and n2 is not nan, t1 and t2 are both relative to the time of the n2th ignition. When t1 and t2 are negative, it indicates before ignition, and when t1 and t2 are positive, it indicates after ignition.

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