On-satellite autonomous task visibility rapid determination method

Abstract

The invention discloses a satellite autonomous task visibility rapid determination method, which comprises the following steps: firstly, according to a given task planning time interval, performing orbit interpolation calculation and orbit discretization at equal intervals, and storing data to avoid the problem of huge data volume caused by continuous calculation; secondly, according to the related state information, the starting moment and the ending moment, visible to the satellite, of the target above the horizon are rapidly determined through a bisection method; and finally, according to the loading capacity of the satellite and the visible starting moment and ending moment of the satellite, performing local search by adopting a dichotomy, and quickly determining the visible starting moment, center moment and ending moment of the satellite load to the target. According to the method, a hierarchical local iteration search strategy is adopted, a calculation result can be rapidly obtained, necessary input information is provided for an autonomous task planning system, planning processing time is shortened, task execution efficiency is improved, and the requirement for high timeliness is met.

Description

Technical Field

[0001] The invention belongs to the technical field of satellite autonomous mission planning, and in particular relates to a method for quickly determining visibility of an on-board autonomous mission. Background Art

[0002] Remote sensing satellites are usually equipped with a variety of payloads such as visible light cameras and infrared cameras, which can provide a large amount of remote sensing satellite data and information. Remote sensing satellites have the characteristics of large-angle rapid maneuverability, and can complete single-strip push-scanning, regional multi-strip mosaic imaging, non-along-track active push-scanning and other functions. According to user mission requirements, the remote sensing satellite control subsystem integrates the on-board autonomous mission planning function. The fundamental goal of autonomous mission planning is to use on-board intelligent technology to autonomously determine the flight mission objectives within the specified time according to the observation mission requirements or real-time guidance information on orbit, in accordance with the satellite resource status and usage constraints, and generate a command sequence to control the satellite's on-orbit operation based on the flight mission objectives.

[0003] According to the principle of orbital dynamics, the satellite flies at a certain orbital angular velocity. For a given mission target, the spatial position of the satellite relative to the target changes with time (e.g. Figure 2 Whether the satellite is visible to the target depends on many factors, such as elevation angle, geocentric angle, relative distance, satellite angle to the target, etc. Therefore, in the initial stage of autonomous mission planning, it is necessary to first determine the time interval within each orbital circle during which the target is visible, and give the start time, center time, and end time of the target visibility to provide input information for subsequent planning. Existing target visibility calculation methods either cannot give the key time set required for mission planning, or are difficult to meet the real-time requirements of calculation under the constraints of onboard computer computing power. Therefore, a method for quickly determining the visibility of onboard targets and autonomous missions under the constraints of onboard computer computing power is needed. Summary of the invention

[0004] The technology of the present invention solves the problem: Overcoming the shortcomings of the prior art, providing a method for quickly determining the visibility of autonomous missions on board a satellite, aiming to meet the needs of autonomous mission planning on board a satellite and the real-time computing requirements under the constraints of the computing power of the onboard computer.

[0005] In order to solve the above technical problems, the present invention discloses a method for quickly determining visibility of autonomous missions on a satellite, comprising:

[0006] Given the task planning time interval and initialize the parameters;

[0007] The given task planning time interval is discretized into equal intervals, and each corresponding time point is recorded as t k ;Use orbit interpolation calculation method to obtain the position vector of the satellite in the earth's fixed coordinate system at each time point;

[0008] According to the position vector of the satellite in the earth's fixed coordinate system at each time point, the time interval of the satellite entering and leaving the horizon in each orbital circle is determined;

[0009] According to the determined time interval of the satellite entering and leaving the horizon in each orbital circle, the moment when the satellite is closest to the target is calculated based on the principle of dichotomy;

[0010] According to the calculated time when the satellite is closest to the target, based on the principle of dichotomy, the earliest visible time of the satellite to the target and the latest visible time of the satellite to the target are calculated.

[0011] In the above-mentioned on-board autonomous mission visibility rapid determination method, the satellite's time interval of entering and leaving the horizon in each orbital circle is determined according to the position vector of the satellite in the earth-fixed coordinate system at each time point, including:

[0012] According to the kth time point t k The satellite's position vector r in the Earth-fixed coordinate system satE (t k ), calculate the time point t k The angle L between the line connecting the satellite and the target and the horizon k ;

[0013] Perform the following iterative calculations:

[0014] 31) According to the k+1th time point t k+1 The satellite's position vector r in the Earth-fixed coordinate system satE (t k+1 ), calculate the time point t k+1 The angle L(t k+1 );

[0015] 32) If L(t k+1 )≥0 and L(t k )<0, then the time t at which the satellite enters the horizon in the current nth orbital circle ns =t k+1 ; Where n represents the number of orbital circles, n=1,2,3,…,N, N represents the number of orbital circles;

[0016] 33) If L(t k+1 )<0 and L(t k )≥0, then the time t at which the satellite leaves the horizon in the current nth orbital circle ne =t k ;

[0017] 34) If the given task planning time interval is not exceeded, return to step 31) to continue iterative calculation;

[0018] 35) If the given mission planning time interval is exceeded, the iteration ends and the number of orbital circles N and the time interval of the satellite entering and leaving the horizon in each orbital circle [t ns ,t ne ].

[0019] In the above-mentioned on-board autonomous mission visibility rapid determination method, according to the determined time interval of the satellite entering and leaving the horizon in each orbital circle, based on the principle of dichotomy, the time when the satellite is closest to the target is calculated, including:

[0020] Determine t ns Time and t ne The angle L(t ns ) and L(t ne );

[0021] Based on the principle of dichotomy, the following iterative calculation is performed:

[0022] 41) Calculate the first central moment

[0023] 42) Determine t m The angle L(t m );

[0024] 43) If L(t ns ) <L(t ne ), then let t ns =t m , L(t ns )=L(t m );

[0025] 44) If L(t ns )≥L(t ne ), then let t ne =t m , L(t es )=L(t m );

[0026] 45) If t ne -t ns If the time is greater than the set time threshold, the process returns to step 41) to continue iterative calculation; otherwise, the iteration ends and the time t′ when the satellite is closest to the target is obtained. m .

[0027] In the above-mentioned on-board autonomous mission visibility rapid determination method, the earliest visible time of the satellite to the target is calculated by the following method:

[0028] According to the calculated time when the satellite is closest to the target, the time interval [t ns,t′ m ];

[0029] In the time interval [t ns ,t′ m ], based on the principle of dichotomy, the following iterations are performed:

[0030] 51) Calculate the second central moment

[0031] 52) Using interpolation method, we can calculate t i The vector r pointing to the satellite at the moment esE (t i ), the position vector r of the satellite in the geocentric fixed coordinate system satE (t i ), and the relative distance d between the satellite and the target es (t i );

[0032] 53) If the calculation determination method selection flag δ is visible Flag = 0, that is, the “geocentric angle” is selected as the visible calculation basis, and the calculated t i The angle θ between the satellite and the target about the center of the earth at the moment GA (t i ), where if θ GA (t i )>θ max or es (t i )>d l , then adjust the starting time and let t ns =t i ; Otherwise, adjust the end time and let t′ m =t i ; where θ max Indicates the angle judgment threshold, d l Represents the relative distance judgment threshold;

[0033] 54) If the calculation determination method selection flag δ is visible Flag =1, that is, "ground elevation angle" is selected as the basis for visibility calculation, then t i The ground elevation angle θ of the satellite relative to the target at time EA (t i ), where if θ EA (t i )>θ max or es (t i )>d l , then adjust the starting time and let t ns =t i ; Otherwise, adjust the end time and let t′m =t i ;

[0034] 55) If the calculation determination method selection flag δ is visible Flag =2, that is, "satellite-to-target angle" is selected as the basis for visibility calculation, then t i The satellite angle to the target at time θ VA (t i ), where if θ VA (t i )>θ max or es (t i )>d l , then adjust the starting time and let t ns =t i ; Otherwise, adjust the end time and let t′ m =t i ;

[0035] 56) If t′ m -t ns If the time is greater than the set time threshold, the process returns to step 51) to continue iterative calculation; otherwise, the iteration ends and the earliest visible time t of the satellite to the target is obtained. see_start .

[0036] In the above-mentioned on-board autonomous mission visibility rapid determination method, the latest visible time of the satellite to the target is calculated as follows:

[0037] According to the calculated time when the satellite is closest to the target, the time interval [t′ m ,t ne ];

[0038] In the time interval [t′ m ,t ne ], based on the principle of dichotomy, the following iterations are performed:

[0039] 61) Calculate the third central moment

[0040] 62) Using interpolation method, we can calculate t j The vector r pointing to the satellite at the moment esE (t j ), the position vector r of the satellite in the geocentric fixed coordinate system satE (t j ), and the relative distance d between the satellite and the target es (t j );

[0041] 63) If the calculation determination method selection flag δ is visible Flag= 0, that is, the “geocentric angle” is selected as the visible calculation basis, and the calculated t j The angle θ between the satellite and the target about the center of the earth at the moment GA (t j ), where if θ GA (t j )>θ max or es (t j )>d l , then adjust the end time, let t ne =t j ; Otherwise, adjust the starting time and let t′ m =t j ;

[0042] 64) If the calculation determination method selection flag δ is visible Flag =1, that is, "ground elevation angle" is selected as the basis for visibility calculation, then t j The ground elevation angle θ of the satellite relative to the target at time EA (t j ), where if θ EA (t j )>θ max or es (t j )>d l , then adjust the end time, let t′ m =t j ; Otherwise, adjust the starting time and let t′ m =t j ;

[0043] 65) If the calculation determination method selection flag δ is visible Flag =2, that is, "satellite-to-target angle" is selected as the basis for visibility calculation, then t j The satellite angle θ to the target at time VA (t j ), where if θ VA (t j )>θ max or es (t j )>d l , then adjust the end time, let t′ m =t j ; Otherwise, adjust the starting time and let t′ m =t j ;

[0044] 66) If t es -t′ mIf the time is greater than the set time threshold, the process returns to step 61) to continue iterative calculation; otherwise, the iteration ends and the latest visible time t of the satellite to the target is obtained. seeend .

[0045] In the above-mentioned on-board autonomous mission visibility rapid determination method, the angle between the line connecting the satellite and the target and the horizon is calculated as follows:

[0046] L(t)=(r satE (t)-r sEp ) T r sEp

[0047] Where L(t) represents the angle between the line connecting the satellite and the target and the horizon at any time t, r satE (t) represents the position vector of the satellite in the Earth-fixed coordinate system at any time t, r sEp The position vector of the target in the earth-fixed coordinate system.

[0048] In the above-mentioned on-board autonomous mission visibility rapid determination method, the calculation formula of the target vector pointing to the satellite is as follows:

[0049] r esE (t) = r satE (t)-r sEp

[0050] Among them, r esE (t) represents the vector pointing from the target to the satellite at any time t.

[0051] In the above-mentioned on-board autonomous mission visibility rapid determination method, the calculation formula of the relative distance between the satellite and the target is as follows:

[0052] d es (t)=||r esE (t)||2

[0053] Among them, d es (t) represents the relative distance between the satellite and the target at any time t.

[0054] The present invention has the following advantages:

[0055] The present invention discloses a method for quickly determining the visibility of an autonomous mission on a satellite. The method discretizes a given mission planning time interval, uses equal spacing to perform orbit interpolation calculation and data storage, and avoids the problem of huge data volume caused by continuous calculation. The binary method is used to quickly determine the start time, center time, and end time of target visibility, and a local iteration strategy is used to quickly obtain calculation results, thereby shortening the planning processing time, improving the mission execution efficiency, and meeting the demand for high timeliness. The method is applicable to visibility judgment of various payloads, judgment based on payload observation range, and rapid visibility judgment based on ground elevation angle missions, and the method has universality. BRIEF DESCRIPTION OF THE DRAWINGS

[0056] Figure 1 is a flow chart of a method for quickly determining visibility of an autonomous mission on board a satellite in an embodiment of the present invention;

[0057] Figure 2 It is a spatial schematic diagram of a satellite relative to a target in an embodiment of the present invention. DETAILED DESCRIPTION

[0058] In order to make the objectives, technical solutions and advantages of the present invention more clear, the embodiments disclosed in the present invention will be further described in detail below with reference to the accompanying drawings.

[0059] One of the core ideas of the present invention is to provide a method for quickly determining the visibility of autonomous tasks on a satellite, taking into account the user's observation task requirements and data transmission requirements, combined with the known state information and usage constraints on the satellite, to quickly determine the visibility and visible time interval of the mission target on orbit. First, according to the given mission planning time interval, the orbit interpolation calculation is performed with equal spacing, the orbit is discretized, and the data is stored to avoid the huge amount of data caused by continuous calculation. Secondly, according to the relevant state information, the binary method is used to quickly determine the start and end time of the target above the horizon to the satellite. Finally, according to the satellite's payload capacity and the satellite's visible start and end time, the binary method is used for local search to quickly determine the satellite payload's visible start time, center time and end time to the target. The method for quickly determining the visibility of autonomous tasks on a satellite proposed by the present invention adopts a hierarchical local iterative search strategy, which can quickly obtain the calculation results, including the visible orbital circles of each mission target, and the key time set in each orbital circle, providing the necessary input information for the autonomous mission planning system. This method shortens the planning processing time, improves the task execution efficiency, and meets the demand for high timeliness. The present invention is universal and can be applied to scenarios including: visibility judgment of various loads, judgment based on load observation range, rapid visibility judgment based on ground elevation angle tasks, etc.

[0060] Reference Figure 1In this embodiment, the method for quickly determining visibility of autonomous missions on a satellite includes:

[0061] Step 1: Give the task planning time interval and initialize the parameters.

[0062] In this embodiment, the parameters that need to be initialized include but are not limited to: Flag , the position vector r of the target in the earth-fixed coordinate system sEp , the target's geographic longitude λ obj , search step dh, angle judgment threshold θ max , relative distance judgment threshold d l wait.

[0063] Step 2: Discretize the given task planning time interval into equal intervals, and record each corresponding time point as t k ; Use the orbit interpolation calculation method to obtain the position vector of the satellite in the earth's fixed coordinate system at each time point.

[0064] In this embodiment, the given mission planning time interval can be discretized into equal intervals according to the injected initial orbit information, and each corresponding time point is recorded as t k ; and use the orbit interpolation calculation method to obtain the position vector of the satellite in the earth's fixed coordinate system at each time point.

[0065] Step 3: Determine the time interval of the satellite entering and leaving the horizon in each orbital cycle based on the position vector of the satellite in the earth's fixed coordinate system at each time point.

[0066] In this embodiment, first, according to the kth time point t k The satellite's position vector r in the Earth-fixed coordinate system satE (t k ), calculate the time point t k The angle L between the line connecting the satellite and the target and the horizon k Then, perform the following iterative calculation:

[0067] 31) According to the k+1th time point t k+1 The satellite's position vector r in the Earth-fixed coordinate system satE (t k+1 ), calculate the time point t k+1 The angle L(t k+1 ).

[0068] 32) If L(t k+1 )≥0 and L(t k )<0, then the time t at which the satellite enters the horizon in the current nth orbital circle ns =tk+1 . Where n represents the number of orbital circles, n=1,2,3,…,N, N represents the number of orbital circles.

[0069] 33) If L(t k+1 )<0 and L(t k )≥0, then the time t at which the satellite leaves the horizon in the current nth orbital circle ne =t k .

[0070] 34) If the given task planning time interval is not exceeded, return to step 31) to continue iterative calculation.

[0071] 35) If the given mission planning time interval is exceeded, the iteration ends and the number of orbital circles N and the time interval of the satellite entering and leaving the horizon in each orbital circle [t ns ,t ne ].

[0072] Among them, the general calculation formula for the angle between the line connecting the satellite and the target and the horizon is as follows:

[0073] L(t)=(r satE (t)-r sEp ) T r sEp …(1)

[0074] Where L(t) represents the angle between the line connecting the satellite and the target and the horizon at any time t, r satE (t) represents the position vector of the satellite in the Earth-fixed coordinate system at any time t.

[0075] Step 4, based on the determined time interval of the satellite entering and leaving the horizon in each orbital circle, based on the principle of dichotomy, calculate the moment when the satellite is closest to the target.

[0076] In this embodiment, first, t ns Time and t ne The angle L(t ns ) and L(t ne ), refer to the above formula (1). Then, based on the principle of dichotomy, perform the following iterative calculation:

[0077] 41) Calculate the first central moment

[0078] 42) Determine t m The angle L(t m ), refer to the above formula (1).

[0079] 43) If L(tns ) <L(t ne ), then let t ns =t m , L(t ns )=L(t m ).

[0080] 44) If L(t ns )≥L(t ne ), then let t ne =t m , L(t es )=L(t m ).

[0081] 45) If t ne -t ns If the time is greater than the set time threshold, the process returns to step 41) to continue iterative calculation; otherwise, the iteration ends and the time t′ when the satellite is closest to the target is obtained. m .

[0082] Step 5, according to the calculated time when the satellite is closest to the target, based on the principle of dichotomy, calculate the earliest visible time and the latest visible time of the satellite to the target.

[0083] In this embodiment, the earliest visible time of the satellite to the target can be calculated as follows:

[0084] According to the calculated time when the satellite is closest to the target, the time interval [t ns ,t′ m ];

[0085] In the time interval [t ns ,t′ m ], based on the principle of dichotomy, the following iterations are performed:

[0086] 51) Calculate the second central moment

[0087] 52) Using interpolation method, we can calculate t i The vector r pointing to the satellite at the moment esE (t i ), the position vector r of the satellite in the geocentric fixed coordinate system satE (t i ), and the relative distance d between the satellite and the target es (t i ).

[0088] 53) If the calculation determination method selection flag δ is visible Flag = 0, that is, the “geocentric angle” is selected as the visible calculation basis, and the calculated t iThe angle θ between the satellite and the target about the center of the earth at the moment GA (t i ). Among them, if θ GA (t i )>θ max or es (t i )>d l , then adjust the starting time and let t ns =t i ; Otherwise, adjust the end time and let t′ m =t i .

[0089] 54) If the calculation determination method selection flag δ is visible Flag =1, that is, "ground elevation angle" is selected as the basis for visibility calculation, then t i The ground elevation angle θ of the satellite relative to the target at time EA (t i ). Among them, if θ EA (t i )>θ max or es (t i )>d l , then adjust the starting time and let t ns =t i ; Otherwise, adjust the end time and let t′ m =t i .

[0090] 55) If the calculation determination method selection flag δ is visible Flag =2, that is, "satellite-to-target angle" is selected as the basis for visibility calculation, then t i The satellite angle θ to the target at time VA (t i ). Among them, if θ VA (t i )>θ max or es (t i )>d l , then adjust the starting time and let t ns =t i ; Otherwise, adjust the end time and let t′ m =t i .

[0091] 56) If t′ m -t ns If the time is greater than the set time threshold, the process returns to step 51) to continue iterative calculation; otherwise, the iteration ends and the earliest visible time t of the satellite to the target is obtained. seestart .

[0092] Furthermore, the latest visible time of the satellite to the target can be calculated as follows:

[0093] According to the calculated time when the satellite is closest to the target, the time interval [t′ m ,t ne ];

[0094] In the time interval [t′ m ,t ne ], based on the principle of dichotomy, the following iterations are performed:

[0095] 61) Calculate the third central moment

[0096] 62) Using interpolation method, we can calculate t j The vector r pointing to the satellite at the moment esE (t j ), the position vector r of the satellite in the geocentric fixed coordinate system satE (t j ), and the relative distance d between the satellite and the target es (t j ).

[0097] 63) If the calculation determination method selection flag δ is visible Flag = 0, that is, the “geocentric angle” is selected as the visible calculation basis, and the calculated t j The angle θ between the satellite and the target about the center of the earth at the moment GA (t j ). Among them, if θ GA (t j )>θ max or es (t j )>d l , then adjust the end time, let t ne =t j ; Otherwise, adjust the starting time and let t′ m =t j .

[0098] 64) If the calculation determination method selection flag δ is visible Flag =1, that is, "ground elevation angle" is selected as the basis for visibility calculation, then t j The ground elevation angle θ of the satellite relative to the target at time EA (t j ). Among them, if θ EA (t j )>θ max or es (t j )>d l , then adjust the end time, let t′m =t j ; Otherwise, adjust the starting time and let t′ m =t j .

[0099] 65) If the calculation determination method selection flag δ is visible Flag =2, that is, "satellite-to-target angle" is selected as the basis for visibility calculation, then t j The satellite angle to the target at time θ VA (t j ). Among them, if θ VA (t j )>θ max or es (t j )>d l , then adjust the end time, let t′ m =t j ; Otherwise, adjust the starting time and let t′ m =t j .

[0100] 66) If t es -t′ m If the time is greater than the set time threshold, the process returns to step 61) to continue iterative calculation; otherwise, the iteration ends and the latest visible time t of the satellite to the target is obtained. see_end .

[0101] Preferably, the general calculation formula for the vector of the target pointing to the satellite is as follows:

[0102] r esE (t) = r satE (t)-r sEp …(2)

[0103] Among them, r esE (t) represents the vector pointing from the target to the satellite at any time t.

[0104] Preferably, the general calculation formula for the relative distance between the satellite and the target is as follows:

[0105] d es (t)=||r esE (t)||2…(3)

[0106] Among them, d es (t) represents the relative distance between the satellite and the target at any time t.

[0107] It should be noted that the geocentric angle θ GA (t), ground elevation angle θ EA (t), satellite angle to target θ VA (t), satellite sub-satellite point geographical longitude λobj (t) can be calculated by interpolation.

[0108] Through the above steps, the earliest visible time, closest position time, and latest visible time of the satellite relative to the target can be obtained for each orbital circle. This time information can provide useful reference input for mission planning.

[0109] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0110] The contents not described in detail in the specification of the present invention belong to the common knowledge of the professionals in this field.

Claims

1. A method for quickly determining visibility of autonomous missions on a satellite, characterized in that: include: Given the task planning time interval and initialize the parameters; The given task planning time interval is discretized into equal intervals, and each corresponding time point is recorded as t k ;Use orbit interpolation calculation method to obtain the position vector of the satellite in the earth's fixed coordinate system at each time point; According to the position vector of the satellite in the earth's fixed coordinate system at each time point, the time interval of the satellite entering and leaving the horizon in each orbital circle is determined; According to the determined time interval of the satellite entering and leaving the horizon in each orbital circle, the moment when the satellite is closest to the target is calculated based on the principle of dichotomy; According to the calculated time when the satellite is closest to the target, based on the principle of dichotomy, the earliest visible time of the satellite to the target and the latest visible time of the satellite to the target are calculated.

2. The method for quickly determining visibility of autonomous missions on board a satellite according to claim 1, characterized in that: Based on the position vector of the satellite in the earth's fixed coordinate system at each time point, the time interval of the satellite entering and leaving the horizon in each orbital circle is determined, including: According to the kth time point t k The satellite's position vector r in the Earth-fixed coordinate system satE (t k ), calculate the time point t k The angle L between the line connecting the satellite and the target and the horizon k ; Perform the following iterative calculations: 31) According to the k+1th time point t k+1 The satellite's position vector r in the Earth-fixed coordinate system satE (t k+1 ), calculate the time point t k+1 The angle L(t k+1 ); 32) If L(t k+1 )≥0 and L(t k )<0, then the time t at which the satellite enters the horizon in the current nth orbital circle ns =t k+1 ; Where n represents the number of orbital circles, n=1,2,3,…,N, N represents the number of orbital circles; 33) If L(t k+1 )<0 and L(t k )≥0, then the time t at which the satellite leaves the horizon in the current nth orbital circle ne =t k ; 34) If the given task planning time interval is not exceeded, return to step 31) to continue iterative calculation; 35) If the given mission planning time interval is exceeded, the iteration ends and the number of orbital circles N and the time interval of the satellite entering and leaving the horizon in each orbital circle [t ns ,t ne ].

3. The method for quickly determining visibility of autonomous missions on board a satellite according to claim 2, characterized in that: According to the determined time interval of the satellite entering and leaving the horizon in each orbital circle, based on the principle of dichotomy, the time when the satellite is closest to the target is calculated, including: Determine t ns Time and t ne The angle L(t ns ) and L(t ne ); Based on the principle of dichotomy, the following iterative calculation is performed: 41) Calculate the first central moment 42) Determine t m The angle L(t m ); 43) If L(t ns ) <L(t ne ), then let t ns =t m , L(t ns )=L(t m ); 44) If L(t ns )≥L(t ne ), then let t ne =t m , L(t es )=L(t m ); 45) If t ne -t ns If the time is greater than the set time threshold, the process returns to step 41) to continue iterative calculation; otherwise, the iteration ends and the time t′ when the satellite is closest to the target is obtained. m .

4. The method for quickly determining visibility of autonomous missions on board a satellite according to claim 3, characterized in that: The earliest visible time of the satellite to the target is calculated as follows: According to the calculated time when the satellite is closest to the target, the time interval [t ns ,t′ m ]; In the time interval [t ns ,t′ m ], based on the principle of dichotomy, the following iterations are performed: 51) Calculate the second central moment 52) Using interpolation method, we can calculate t i The vector r pointing to the satellite at the moment esE (t i ), the position vector r of the satellite in the geocentric fixed coordinate system satE (t i ), and the relative distance d between the satellite and the target es (t i ); 53) If the calculation determination method selection flag δ is visible Flag = 0, that is, "Earth Center Angle" is selected as the visible calculation basis, then the calculated t i The angle θ between the satellite and the target about the center of the earth at the moment GA (t i ), where if θ GA (t i )>θ max or es (t i )>d l , then adjust the starting time and let t ns =t i ; Otherwise, adjust the end time and let t′ m =t i ; where θ max Indicates the angle judgment threshold, d l Represents the relative distance judgment threshold; 54) If the calculation determination method selection flag δ is visible Flag =1, that is, "ground elevation angle" is selected as the basis for visibility calculation, then t i The ground elevation angle θ of the satellite relative to the target at time EA (t i ), where if θ EA (t i )>θ max or es (t i )>d l , then adjust the starting time and let t ns =t i ; Otherwise, adjust the end time and let t′ m =t i ; 55) If the calculation determination method selection flag δ is visible Flag =2, that is, "satellite-to-target angle" is selected as the basis for visibility calculation, then t i The satellite angle θ to the target at time VA (t i ), where if θ VA (t i )>θ max or es (t i )>d l , then adjust the starting time and let t ns =t i ; Otherwise, adjust the end time and let t′ m =t i ; 56) If t′ m -t ns If the time is greater than the set time threshold, the process returns to step 51) to continue iterative calculation; otherwise, the iteration ends and the earliest visible time t of the satellite to the target is obtained. see_start .

5. The method for quickly determining visibility of autonomous missions on board a satellite according to claim 4, characterized in that: The latest visible time of the satellite to the target is calculated as follows: According to the calculated time when the satellite is closest to the target, the time interval [t′ m ,t ne ]; In the time interval [t′ m ,t ne ], based on the principle of dichotomy, the following iterations are performed: 61) Calculate the third central moment 62) Using interpolation method, we can calculate t j The vector r pointing to the satellite at the moment esE (t j ), the position vector r of the satellite in the geocentric fixed coordinate system satE (t j ), and the relative distance d between the satellite and the target es (t j ); 63) If the calculation determination method selection flag δ is visible Flag = 0, that is, "Earth Center Angle" is selected as the visible calculation basis, then the calculated t j The angle θ between the satellite and the target about the center of the earth at the moment GA (t j ), where if θ GA (t j )>θ max or es (t j )>d l , then adjust the end time, let t ne =t j ; Otherwise, adjust the starting time and let t′ m =t j ; 64) If the calculation determination method selection flag δ is visible Flag =1, that is, "ground elevation angle" is selected as the basis for visibility calculation, then t j The ground elevation angle θ of the satellite relative to the target at time EA (t j ), where if θ EA (t j )>θ max or es (t j )>d l , then adjust the end time, let t′ m =t j ; Otherwise, adjust the starting time and let t′ m =t j ; 65) If the calculation determination method selection flag δ is visible Flag =2, that is, "satellite-to-target angle" is selected as the basis for visibility calculation, then t j The satellite angle θ to the target at time VA (t j ), where if θ VA (t j )>θ max or es (t j )>d l , then adjust the end time, let t′ m =t j ; Otherwise, adjust the starting time and let t′ m =t j ; 66) If t es -t′ m If the time is greater than the set time threshold, the process returns to step 61) to continue iterative calculation; otherwise, the iteration ends and the latest visible time t of the satellite to the target is obtained. see_end .

6. The method for quickly determining visibility of autonomous missions on board a satellite according to claim 5, characterized in that: The angle between the line connecting the satellite and the target and the horizon is calculated as follows: L(t)=(r satE (t)-r sEp ) T r sEp Where L(t) represents the angle between the line connecting the satellite and the target and the horizon at any time t, r satE (t) represents the position vector of the satellite in the Earth-fixed coordinate system at any time t, r sEp The position vector of the target in the earth-fixed coordinate system.

7. The method for quickly determining visibility of autonomous missions on board a satellite according to claim 6, characterized in that: The calculation formula of the target pointing to the satellite vector is as follows: r esE (t)=r satE (t)-r sEp Among them, r esE (t) represents the vector pointing from the target to the satellite at any time t.

8. The method for quickly determining visibility of autonomous missions on board a satellite according to claim 7, characterized in that: The calculation formula of the relative distance between the satellite and the target is as follows: d es (t)=||r esE (t)||2 Among them, d es (t) represents the relative distance between the satellite and the target at any time t.

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