A method for autonomous mission planning on satellite based on orbital solar vector

Abstract

The invention discloses a method for realizing autonomous mission planning on a satellite based on the solar vector of an orbital system. A ground mission planning system plans the mission and determines the solar altitude angle information at the mission observation time. The operation control system plans the instruction sequence according to the information provided by the ground mission planning system. The instruction sequence includes the mission mode word, the load parameter configuration, and the solar altitude angle modification instruction at the observation time, and the instruction sequence does not include the guidance law of the platform and the turntable and the load on / off instruction. When the satellite enters the country, the instruction sequence is added. The on-board software enters the corresponding mission according to the mission mode word, calculates the key points of the mission start time and end time, and autonomously executes the observation mission. The observation data is transmitted to the ground mission planning system through digital transmission. The method can reduce the number of instructions in the instruction sequence added on the ground, reduce the operating burden of ground personnel, reduce the observation error caused by the delay of the ground addition, and reduce the probability of making mistakes.

Description

Technical Field

[0001] The invention belongs to the field of satellite autonomous mission planning, and provides a method for autonomous mission planning of a sun-synchronous near-circular orbit satellite, specifically a method for realizing on-board autonomous mission planning based on the orbital system sun vector. Background Art

[0002] Satellite mission planning is part of resource management from the perspective of systematic application. At present, the execution of satellite on-orbit missions is mainly based on the satellite on-orbit energy status, attitude information, payload unit health status and orbit information received by the ground measurement and control system. The ground operator negotiates with the satellite user to select the observation target and time, and the ground measurement and control system injects the instruction sequence to achieve the goal. The instruction sequence is a set of instructions, which usually includes instructions for controlling the timing of task execution, payload unit switch instructions, and payload unit parameter configuration instructions. In order to achieve observation accuracy and make full use of space resources, a single satellite carries more and more payloads. Mission observation is achieved through the cooperation between various payloads. If all mission observations are achieved by injecting instruction sequences on the ground, the instructions that need to be injected on the ground will inevitably become more and more complex. The production of instruction sequences is time-consuming, labor-intensive and error-prone. If the satellite's entry time is short and the instruction sequence cannot be injected in time, it may result in the observation mission not being completed in time or the observation time being reduced.

[0003] At present, the processor performance of the onboard integrated electronic system has reached more than 200MHz, and can even reach 1GHz using screened industrial-grade processors. Combined with the large memory capacity, it is fully capable of transferring some ground mission planning functions to onboard processing.

[0004] Purpose of the Invention

[0005] In view of the problems that the current satellite observation mission mainly relies on the command sequence injected on the ground and the ground operation is complicated and prone to errors caused by controlling the onboard payload to perform observation alone, the present invention proposes a method for planning tasks for sun-synchronous near-circular orbit satellites based on the orbital system sun vector. The method can reduce the number of instructions in the command sequence injected on the ground, alleviate the operating burden of ground personnel, reduce the observation error caused by the delay of ground injection, and reduce the probability of error.

[0006] The technical solution of the present invention is: a method for realizing autonomous mission planning on a satellite based on the orbital system solar vector, and the specific steps are as follows:

[0007] Step 1: The ground mission planning system plans the mission and determines the solar altitude angle information at the mission observation time;

[0008] Step 2: The operation control system plans the instruction sequence according to the information provided by the ground mission planning system, wherein the instruction sequence includes the mission mode word, the load parameter configuration, and the solar altitude angle modification instruction at the observation time, and the instruction sequence does not include the guidance law of the platform and the turntable and the load power on / off instruction;

[0009] Step 3: When the satellite enters the country, the command sequence is entered;

[0010] Step 4: The onboard software enters the corresponding task according to the task mode word, calculates the key points of the task start and end time, and autonomously performs the observation task;

[0011] Step 5: Transmit the observation data to the ground mission planning system via digital transmission.

[0012] Furthermore, in step 4, the onboard software enters the corresponding task according to the task mode word, calculates the key points of the task start time and end time, and autonomously performs the observation task; the details are as follows:

[0013] For a sun-synchronous circular orbit satellite, the Z axis of the satellite mass center orbit coordinate system is defined as the line connecting the satellite to the Earth's mass center, pointing to the Earth's mass center; the Y axis of the satellite mass center orbit coordinate system is in the same direction as the negative normal of the orbital plane; the X axis of the satellite mass center orbit coordinate system forms a right-handed orthogonal system with the Y and Z axes, pointing in the forward direction in the orbital plane; the satellite mass center orbit coordinate system is the VVLH coordinate system;

[0014] The satellite attitude control module can derive the normalized orbital solar vector μso based on the solar ephemeris model and orbital data; the various angles of the solar vector μso in the VVLH system are defined as follows: the φ angle is the angle between the projection of the solar vector μso on the orbital plane and the z-axis, which is positive around the Y-axis; the α angle is the angle between the solar vector μso and its projection on the xoy plane, that is, the solar altitude angle, which is negative toward the Z-axis; the β angle is the angle between the solar vector μso and its projection on the orbital plane, that is, the orbital solar angle, which is negative toward the Y-axis; the θ angle is the angle between the projection of the solar vector μso and the x-axis, which is positive around the Z-axis;

[0015] The relationship between the sun vector and each angle is derived as follows:

[0016]

[0017] The calculation formulas for α, β, θ, and φ can be obtained from formula (1):

[0018]

[0019] α=-asin(μso(3)) (3)

[0020] β=asin(μso(2)) (4)

[0021] θ=atan2(μso(2),μso(1)) (5)

[0022] According to the orbital characteristics, in the same orbit observation, the φ angle changes linearly within [-π,π], and the β angle remains almost unchanged. The relationship between the φ angle and the α and β angles is as shown in formulas (6) and (7);

[0023] Sun Rising Phase:

[0024]

[0025] Sun setting phase:

[0026]

[0027] The formula for calculating α from φ and β is as follows:

[0028]

[0029] The calculation formula of angle θ and α and β is as follows:

[0030] Sun Rising Segment:

[0031]

[0032] Sun setting section:

[0033]

[0034] If the value of the solar altitude angle α at the mission observation time designed by the ground mission planning system is known, the value of φ at the observation time can be calculated from α and β. Since the value of φ changes linearly, the on-board time t_flag(i) (i = 1, 2, 3...) at which the mission observation starts and ends can be derived, as shown in formula (11). The on-board software calculates t_flag(i) at regular intervals to eliminate errors. When the on-board time is equal to t_flag(i), the autonomous task management module of the on-board software outputs corresponding instructions and executes related operations to start or end mission observation according to the current satellite energy situation, satellite attitude information, orbit information, and the health status of the payload unit.

[0035]

[0036] Where: t_flag(i) is the start and end time points on the satellite of the mission observation;

[0037] is the φ value corresponding to the solar altitude angle at the beginning of observation;

[0038] is the φ value corresponding to the current satellite position;

[0039] T_sat is the current time on the satellite;

[0040] ω is the spacecraft orbital angular velocity, which is given by the orbit module.

[0041] Furthermore, the satellite center of mass orbit coordinate system X-axis is defined to form a right-handed orthogonal system with Y-axis and Z-axis, pointing to the forward direction in the orbital plane, and if it is a circular orbit, it is consistent with the velocity direction.

[0042] The beneficial effects of the present invention are as follows: unlike traditional satellite mission planning, the instruction sequence does not require planning of platform and turntable guidance laws, load power on and off instructions, etc., which effectively reduces the number of instructions to be uploaded and improves the success rate of the upload.

[0043] The method of planning satellite missions using the orbital solar vector calculates the relative position and angle between the satellite and the sun through the solar vector, and plans the execution timing of the satellite mission according to the time when the satellite reaches the specified angle. It can discretize the visible observation time window and execute different observation tasks in different time periods, thus realizing the mutual exclusion and autonomous execution of tasks. It has the advantages of saving measurement and control channel resources, reducing human intervention, and improving the efficiency of on-board resource utilization. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] Figure 1 This is a schematic diagram of the angle of the sun vector in the orbital coordinate system;

[0045] Figure 2 This is a schematic diagram of the corresponding relationship between the solar altitude angle and the observation interval;

[0046] Figure 3 This is a schematic diagram of the onboard autonomous mission process. DETAILED DESCRIPTION

[0047] The present invention will be further described below in conjunction with the accompanying drawings.

[0048] For a sun-synchronous circular orbit satellite, the satellite mass center orbit coordinate system (VVLH coordinate system) is defined as the Z axis of the satellite mass center orbit coordinate system (VVLH coordinate system) as the line connecting the satellite to the Earth's mass center, pointing to the Earth's mass center; the Y axis is in the same direction as the negative normal of the orbital plane, and the X axis forms a right-handed orthogonal system with the Y and Z axes, pointing in the forward direction in the orbital plane (in the case of a circular orbit, it is consistent with the velocity direction). The satellite attitude control module can derive the normalized orbital system solar vector μso based on the solar ephemeris model and orbital data. Then the various angles of the solar vector μso in the VVLH system are defined as follows Figure 1As shown: where the angle φ is the angle between the projection of the solar vector μso on the orbital plane and the z-axis, which is positive around the Y-axis; the angle α is the angle between the solar vector μso and its projection on the xoy plane (i.e. the solar altitude angle), which is negative toward the Z-axis; the angle β is the angle between the solar vector μso and its projection on the orbital plane (i.e. the orbital solar angle), which is negative toward the Y-axis; the angle θ is the angle between the projection of the solar vector μso and the x-axis, which is positive around the Z-axis.

[0049] The relationship between the sun vector and each angle is derived as follows:

[0050]

[0051] The calculation formulas for α, β, θ, and φ can be obtained from formula (1).

[0052]

[0053] α=-asin(μso(3)) (3)

[0054] β=asin(μso(2)) (4)

[0055] θ=atan2(μso(2),μso(1)) (5)

[0056] According to the orbital characteristics, in the same orbit observation, the φ angle changes linearly in [-π,π], and the β angle remains almost unchanged. The relationship between the φ angle and the α and β angles is shown in formulas (6) and (7).

[0057] Sun Rising Phase:

[0058]

[0059] Sun setting phase:

[0060]

[0061] The formula for calculating α from φ and β is as follows:

[0062]

[0063] The calculation formula of angle θ and α and β is as follows:

[0064] Sun Rising Segment:

[0065]

[0066] Sun setting section:

[0067]

[0068] If the value of the solar altitude angle α at the mission observation time designed by the ground mission planning system is known (it can be modified by command annotation), the value of φ at the observation time can be calculated from α and β. Since the value of φ changes linearly, the on-board time t_flag(i) (i = 1, 2, 3...) at which the mission starts and ends can be derived, as shown in formula (11). The on-board software calculates t_flag(i) at regular intervals to eliminate errors. When the on-board time is equal to t_flag(i), the autonomous task management module of the on-board software outputs corresponding instructions based on the current satellite energy situation, satellite attitude, orbit information, and payload stand-alone health status, and executes related operations to start or end mission observation.

[0069]

[0070] Where: t_flag(i) is the start and end time points on the satellite of the mission observation;

[0071] is the φ value corresponding to the solar altitude angle at the beginning of observation;

[0072] is the φ value corresponding to the current satellite position;

[0073] T_sat is the current time on the satellite;

[0074] ω is the spacecraft orbital angular velocity, which is given by the orbit module.

[0075] The execution process of the whole-star mission observation is as follows:

[0076] The ground mission planning system plans the mission and determines the solar altitude angle information at the mission observation time;

[0077] The operation control system plans the instruction sequence according to the information provided by the ground mission planning system, and the instruction sequence includes the mission mode word, the load parameter configuration, and the solar altitude angle modification instruction at the observation time, and the instruction sequence does not include the guidance law of the platform and turntable and the load power on and off instructions.

[0078] When the satellite enters, the command sequence is added;

[0079] The onboard software enters the corresponding task according to the mission mode word, calculates the key points of the mission start and end time, and autonomously performs the observation mission.

[0080] The observation data is transmitted to the ground mission planning system via digital transmission.

[0081] The above operation is different from traditional satellite mission planning in that the instruction sequence does not require planning of platform and turntable guidance laws, payload power on and off instructions, etc., which effectively reduces the number of instructions to be uploaded and improves the success rate of upload.

[0082] At present, satellites carry more and more payloads, and their missions are becoming more and more complex. Due to the limitations of space resources and the environment, satellite attitude adjustment and payload power-on and power-off operations need to be performed according to observation missions. Aiming at the characteristics of low-orbit sun-synchronous near-circular orbit satellites, the present invention proposes a satellite autonomous mission planning method based on the orbital system sun vector, which performs on-board autonomous calculations of subsequent observation times based on the mission start time noted on the ground, and plans payload power-on and power-off operations and the sending of guidance laws based on the calculation results. The following is an example to illustrate the specific implementation of the present invention:

[0083] The ground user designs the observation sun altitude angle in advance and plans the start time of the mission observation. The planned sun altitude angle sequence in the example of the present invention is sun rise -17°, sun rise -13°, sun rise -5°, sun rise 5°, sun rise 20° for about 75 seconds, sun altitude angle maximum αmax for about 75 seconds, sun set 20° for about 75 seconds, sun set 5°, sun set -5°, sun set -13°, sun set -17°. There are 14 points in total.

[0084] The mission start time on the ground is t0. When the satellite receives the mission start information, it starts to calculate the time on the satellite when the observation angle is in place, which is t1 to t14. The corresponding relationship between the solar altitude angle and the observation interval is as follows: Figure 2 shown.

[0085] Compare the interval between t1 and t0. Since the payload needs to be turned on 600 seconds in advance, if (t1-t0) < 600, the mission on orbit is abandoned and the time when the sun rises next time and the solar altitude angle is -17° is calculated. In order to reduce the error, the satellite repeatedly calculates t1 to t14 at a certain interval. The satellite software determines the size of the current on-board time and t1 to t14 and performs the following operations:

[0086] 1.t1-600s: Send the power-on command for load unit 1.

[0087] 2.t1-591s: Configure the parameters of load unit 1 according to the single unit parameter table noted on the ground.

[0088] 3.t1-240s: Enter the solar observation sub-mode and send the platform tracking guidance law.

[0089] 4.t1~t2: The payload performs the solar observation mission.

[0090] 5.t2: End solar observation, send the platform to erect the ground guidance law, shut down payload unit 1, and switch to the 0° observation mission in the sub-mode.

[0091] 6.t3-120s: Send power-on commands to load units 2 and 3 at intervals of 1s.

[0092] 7.t3-110s: Configure load unit 2 and 3 parameters according to the unit parameter table noted on the ground.

[0093] 8.t3-63s: Turntable on.

[0094] 9.t3-60s: Send power-on instructions for loads 4 and 5 in sequence.

[0095] 10.t3-50s: Configure load unit 4 and 5 parameters according to the unit parameter table noted on the ground.

[0096] 11.t3-40s: The turntable moves to the predetermined position (determined by the angle θ) and is ready for observation.

[0097] 12. t3~t4: Start 0° mission observation. During the observation, the turntable will position and observe for 1 second before moving to the next position, and repeat this cycle until the observation is completed.

[0098] 13.t4: The platform is erected and the turntable is positioned, ending the 0° observation and setting the sub-mode word to 20° observation.

[0099] 14.t4+2s: The turntable maneuvers to the next observation position.

[0100] 15.t5-40s: Platform tracking, preparing for 20° observation.

[0101] 16.t5~t6: Platform tracking, start 20° observation mission.

[0102] 17.t6: End the 20° observation, erect the platform, shut down the turntable, turn off loads 2, 3, 4, and 5 at a time interval of 1s, and set the observation sub-mode word to αmax observation.

[0103] 18.t7-63s: Turntable on.

[0104] 19.t7-53s: The turntable moves to the predetermined position.

[0105] 20.t7-40s: Turntable positioning, platform tracking, prepare for observation.

[0106] 21.t7~t8: Start αmax observation task.

[0107] 22.t8: After the observation is finished, the platform is erected, the turntable returns to its original position and then the machine is turned off.

[0108] twenty three.……

[0109] 24.t14: End of solar observation mission, all payloads shut down, platform erected, and turntable shut down.

[0110] The above steps only describe the autonomous execution process of the mission during the rising phase of the sun. The landing phase is roughly the same as the rising phase, and the specific operation process also involves communication with the payload, telemetry and remote control processing, etc., which will not be repeated here. Figure 3 shown.

[0111] The method of planning satellite missions using the orbital solar vector calculates the relative position and angle between the satellite and the sun through the solar vector, and plans the execution timing of the satellite mission according to the time when the satellite reaches the specified angle. It can discretize the visible observation time window and execute different observation tasks in different time periods, thus realizing the mutual exclusion and autonomous execution of tasks. It has the advantages of saving measurement and control channel resources, reducing human intervention, and improving the efficiency of on-board resource utilization.

[0112] The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principle of the present invention. These improvements and modifications should also be regarded as the scope of protection of the present invention.

Claims

1. A method for realizing autonomous mission planning on a satellite based on the orbital system solar vector, characterized in that: The specific steps are as follows: Step 1: The ground mission planning system plans the mission and determines the solar altitude angle information at the mission observation time; Step 2: The operation control system plans the instruction sequence according to the information provided by the ground mission planning system, wherein the instruction sequence includes the mission mode word, the load parameter configuration, and the solar altitude angle modification instruction at the observation time, and the instruction sequence does not include the guidance law of the platform and the turntable and the load power on / off instruction; Step 3: When the satellite enters the country, the command sequence is entered; Step 4: The onboard software enters the corresponding task according to the task mode word, calculates the key points of the task start and end time, and autonomously performs the observation task; Step 5: Transmit the observation data to the ground mission planning system via digital transmission.

2. The method for realizing autonomous mission planning on a satellite based on the orbital solar vector according to claim 1, characterized in that: In step 4, the onboard software enters the corresponding task according to the task mode word, calculates the key points of the task start and end time, and autonomously performs the observation task; the details are as follows: For a sun-synchronous circular orbit satellite, the Z axis of the satellite mass center orbit coordinate system is defined as the line connecting the satellite to the Earth's mass center, pointing to the Earth's mass center; The negative normal direction of the Y-axis orbital plane of the satellite's center of mass orbital coordinate system is consistent; The X axis of the satellite mass center orbit coordinate system forms a right-handed orthogonal system with the Y axis and the Z axis, and points to the forward direction in the orbital plane; the satellite mass center orbit coordinate system is the VVLH coordinate system; The satellite attitude control module can derive the normalized orbital solar vector μso based on the solar ephemeris model and orbital data; the various angles of the solar vector μso in the VVLH system are defined as follows: the φ angle is the angle between the projection of the solar vector μso on the orbital plane and the z-axis, which is positive around the Y-axis; the α angle is the angle between the solar vector μso and its projection on the xoy plane, that is, the solar altitude angle, which is negative toward the Z-axis; the β angle is the angle between the solar vector μso and its projection on the orbital plane, that is, the orbital solar angle, which is negative toward the Y-axis; the θ angle is the angle between the projection of the solar vector μso and the x-axis, which is positive around the Z-axis; The relationship between the sun vector and each angle is derived as follows: The calculation formulas for α, β, θ, and φ can be obtained from formula (1): α= -asin (μso(3)) (3) β=asin(μso(2)) (4) θ=atan2(μso(2),μso(1)) (5) According to the orbital characteristics, in the same orbit observation, the φ angle changes linearly within [-π,π], and the β angle remains almost unchanged. The relationship between the φ angle and the α and β angles is as shown in formulas (6) and (7); Sun Rising Phase: Sun setting phase: The formula for calculating α from φ and β is as follows: The calculation formula of angle θ and α and β is as follows: Sun Rising Segment: Sun setting section: If the value of the solar altitude angle α at the mission observation time designed by the ground mission planning system is known, the value of φ at the observation time can be calculated from α and β. Since the value of φ changes linearly, the on-board time t_flag(i) (i = 1, 2, 3...) at which the mission observation starts and ends can be derived, as shown in formula (11). The on-board software calculates t_flag(i) at regular intervals to eliminate errors. When the on-board time is equal to t_flag(i), the autonomous task management module of the on-board software outputs corresponding instructions and executes related operations to start or end mission observation according to the current satellite energy situation, satellite attitude information, orbit information, and the health status of the payload unit. Where: t_flag(i) is the start and end time points on the satellite of the mission observation; is the φ value corresponding to the solar altitude angle at the beginning of observation; is the φ value corresponding to the current satellite position; Tsat is the current time on the satellite; ω is the spacecraft orbital angular velocity, which is given by the orbit module.

3. The method for realizing autonomous mission planning on a satellite based on the orbital solar vector according to claim 2, characterized in that: The satellite mass center orbit coordinate system X-axis is defined as a right-handed orthogonal system with Y and Z axes, pointing in the forward direction in the orbital plane, and if it is a circular orbit, it is consistent with the velocity direction.

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