A method and device for autonomously planning a shadow forecast of a near-earth remote sensing satellite

Abstract

A kind of near-earth remote sensing satellite autonomous mission planning umbra prediction method and device, based on the relative position relationship of sun-earth-satellite and cylindrical shadow model, first, the time of satellite's out of shadow and the time of satellite's into shadow in each orbit are estimated;Then, in the estimated value of the time of satellite's out of shadow and the time of satellite's into shadow, respectively, the nonlinear equation that the time of satellite's out of shadow and the time of satellite's into shadow satisfy is solved using bisection method iteration, to obtain the accurate result of the time of satellite's into shadow.The present application does not contain complex trigonometric function operation, and the number of iteration solving is less, can satisfy the high real-time requirement and time accuracy requirement of on-board autonomous mission planning, has very high engineering application value.

Description

Technical Field

[0001] This invention relates to the field of spacecraft autonomous mission planning technology, and in particular to a method and apparatus for predicting the Earth's shadow during autonomous mission planning of near-Earth remote sensing satellites. This method is used to calculate the entry and exit times of the Earth's shadow for each orbit of the satellite within a specified planning period when autonomously planning and arranging flight missions on board. Background Technology

[0002] Compared to traditional satellite operation and management that relies primarily on remote control and telemetry, on-board autonomous mission planning offers numerous advantages, including fully leveraging satellite utilization efficiency, enhancing satellite's ability to respond to emergency missions, significantly reducing reliance on ground-based telemetry and control, and supporting multi-satellite autonomous collaborative Earth observation missions. It will become a standard technology for intelligent satellites / constellations. In the autonomous mission planning process, forecasting the time when the satellite enters and exits the Earth's shadow for each orbital circle within the specified planning period is an essential and crucial step before mission scheduling and arrangement. This is because: (1) the satellite usually switches its flight state (from solar to Earth or from Earth to solar) at the moment of entering / exiting the shadow, and mission planning needs to avoid scheduling observation or playback missions near the time of entering or exiting the shadow; (2) the working conditions of some payloads are affected by the Earth's shadow, for example, visible light cameras can generally only image in the sunlit area; (3) the attitude maneuvering actions of the satellite are different when performing missions inside and outside the Earth's shadow area. For example, when the observation mission is completed and there is sufficient idle time afterward, the satellite in the sunlit area should switch to solar charging, and the satellite in the shadow area should switch to zero attitude towards the Earth; when performing playback missions, the satellite in the sunlit area needs to switch to zero attitude towards the Earth, while the satellite in the shadow area does not need attitude maneuvering. It can be seen that accurate Earth shadow forecasting is the basis for the correct result of mission planning and arrangement; at the same time, autonomous mission planning has high real-time requirements, and the faster the Earth shadow forecast is executed, the better.

[0003] Due to the limited computing and storage capabilities of onboard processors, the traditional method of solving shadow conditions using a fixed step size is unsuitable due to the excessive computational load. Jia Xianghua et al. proposed an onboard algorithm for shadow prediction of near-Earth orbit satellites (Jia Xianghua, Xu Ming, Chen Luojing. Shadow Prediction Algorithm for Near-Earth Orbit Satellites [J], Journal of Astronautics, 2016, Vol. 37, No. 1). The algorithm has a relatively small computational load, but its shortcomings for practical engineering applications are as follows: (1) It involves a large number of trigonometric function operations in both iterative calculation and analytical solution, and the computational load is still too large. Existing onboard processors will take more than 20 seconds to predict the shadow time of a day using this algorithm, which cannot meet the higher real-time requirements of autonomous mission planning; (2) The variance of its prediction results is large and not directly related to the prediction duration, and the prediction accuracy needs to be further improved. Summary of the Invention

[0004] The technical problem solved by this invention is that the computational timeliness and accuracy of the existing satellite shadow prediction algorithm cannot meet the needs of satellite autonomous mission planning. Therefore, this invention proposes a method and device for shadow prediction in near-Earth remote sensing satellite autonomous mission planning.

[0005] The technical solution of the present invention is: a method for predicting the shadow of a near-Earth remote sensing satellite autonomous mission planning, comprising: (1) determining the planning period of the autonomous mission planning as [T0, T1], and taking the start time T0 of the mission planning period as the initialization time t0; (2) calculating the estimated shadow time t of the satellite in the orbital sphere to which the satellite belongs at time t0 based on the relative positional relationship between the Sun, Earth and the satellite. D_est And the estimated value of the entry time t C_est (3) Based on the estimated value t of the emergence time D_est And the estimated value of the entry time t C_est The precise value t of the shadow time is obtained by iteratively solving using the bisection method. D And the precise value of the projection time t C (4) The precise value of the advance time t C Maximum length T of satellite shadow area Umax The sum plus the duration margin q is used as the final initialization time t0; (5) Determine whether the final initialization time t0 is less than the end time T1 of the task planning cycle. If so, go to step (2) and continue to calculate the entry and exit time of the next orbital circle; otherwise, the entry and exit times of all orbital circles corresponding to the autonomous task planning cycle are calculated.

[0006] Furthermore, in step (2), the estimated emergence time t of the satellite in the orbital sphere to which t0 belongs is calculated. D_est And the estimated value of the entry time t C_est Specifically, it includes:

[0007] (2a) Determine the instantaneous orbital elements of the satellite at time t0, including the semi-major axis a0, the right ascension of the intersection Ω0, the orbital inclination i0, the perigee argument ω0 and the true perigee argument f0, the latitude argument u0=ω0+f0, and calculate the inertial position vector r0 of the satellite at time t0;

[0008] (2b) Calculate the unit vector s of the sun's direction in the geocentric inertial frame at time t0 by calculating the orbital elements of the sun in the inertial frame, and then transform it to the orbital coordinate system O-ξηζ to obtain the result.

[0009]

[0010] In this system, the origin of the orbital coordinate system is the geocenter O, the ξ-axis coincides with the radius vector of the ascending node N, the ζ-axis coincides with the orbital angular momentum vector, the η-axis is determined by the right-hand rule, and the ξη plane is the orbital plane.

[0011] (2c) Based on Calculate the cosine of the angle β between s and the orbital plane:

[0012]

[0013] in, and They are Components along the ξ and η axes;

[0014] (2d) Calculate the angle between the satellite position vector r and s at the moment of shadow point. Approximate value of cosine:

[0015]

[0016] Where R is the average radius of the Earth;

[0017] (2e) Based on cosβ and Calculate the estimated angle u1 between the projection of s onto the orbital plane and OD using the spherical trigonometric formula, where point D is the exit point of the shadow, and π / 2 < u1 < π.

[0018]

[0019] (2f) Calculate the angle α between the projection of s onto the track plane and ON, where 0 ≤ α < 2π.

[0020]

[0021] (2g) Calculate the estimated time of shadow t based on u1 and α respectively. D_est And the estimated value of the entry time t C_est :

[0022] t D_est =t+(α-u1-u0) / n,

[0023] t C_est =t+(α+u1-u0) / n,

[0024] Among them, satellite orbital angular velocity μ=398600.44km 3 / s 2 It is the gravitational constant.

[0025] Furthermore, in step (3), the precise value t of the shadow time is obtained by iteratively solving the problem using the bisection method. D The precise value t of the projection time C Specifically, it includes:

[0026] (3a) Based on the relative positions of the Sun, Earth, and satellite at the entrance point C and exit point D, we obtain:

[0027]

[0028] Establish t D and t C The nonlinear solution equation is: g(t) = 0;

[0029] in, t represents any moment in the satellite's orbital flight, r(t) is the position vector of the satellite in the geocentric inertial coordinate system at moment t obtained by analytical extrapolation, and r(t) is the length of r(t);

[0030] (3b) Set t D The solution interval is [t] D_est -p,t D_est +p],t C The solution interval is [t] C_est -p,t C_est +p], the upper and lower bounds of the two intervals are uniformly represented by [t]. low , t up ]express;

[0031] (3c) Calculate t mid =(t up +t low ) / 2;

[0032] (3d) If t up -t low ≤ε,t mid As an optimization result, the emergence time t D =t mid , the moment of entering the film C =t mid Stop the calculation; otherwise, if g(t) low )g(t mid If t < 0, let t up =t mid If g(t) low )g(t mid If t > 0, let t low =t mid Then return to step (3c) to proceed to the next iteration; where ε represents the level of precision.

[0033] Furthermore, the accuracy level ε = 0.5s.

[0034] Furthermore, for orbits of 500–700 km, the maximum duration T of the satellite shadow region is... Umax Take 2400s.

[0035] A near-Earth remote sensing satellite autonomous mission planning and shadow prediction device includes: an initialization module, a shadow emergence and arrival time estimation module, a shadow emergence and arrival time determination module, an initialization time determination module, and a mission planning module. The initialization module determines the autonomous mission planning period as [T0, T1], using the start time T0 of the mission planning period as the initialization time t0. The shadow emergence and arrival time estimation module calculates the estimated shadow emergence time t0 of the satellite in its orbital sphere at time t0 based on the relative positional relationship between the Sun, Earth, and the satellite. D_est And the estimated value of the entry time t C_est The shadow exit and entry time determination module is based on the estimated shadow exit time t. D_est And the estimated value of the entry time t C_est The precise value t of the shadow time is obtained by iteratively solving using the bisection method. D And the precise value of the projection time t C The initialization timing determination module determines the precise value t of the emergence time. C Maximum length T of satellite shadow area Umax The sum of these values ​​plus a duration margin q is used as the final initialization time t0. The task planning module determines whether the final initialization time t0 is less than the termination time T1 of the task planning cycle. If so, the exit and entry time estimation module, the exit and entry time determination module, and the initialization time determination module continue to calculate the exit and entry time of the next orbital circle. Otherwise, the exit and entry times of all orbital circles corresponding to the autonomous task planning cycle have been calculated.

[0036] Furthermore, the emergence and emergence time estimation module calculates the predicted emergence time t of the satellite in the orbital sphere to which t0 belongs. D_est And the predicted time of advance t C_est Specifically, it includes:

[0037] (2a) Determine the instantaneous orbital elements of the satellite at time t0, including the semi-major axis a0, the right ascension of the intersection Ω0, the orbital inclination i0, the perigee argument ω0 and the true perigee argument f0, the latitude argument u0=ω0+f0, and calculate the inertial position vector r0 of the satellite at time t0;

[0038] (2b) The unit vector s of the sun's direction in the geocentric inertial frame at time t0 is calculated by calculating the orbital elements of the sun in the inertial frame, and then transformed into the orbital coordinate system O-ξηζ to obtain the following result.

[0039]

[0040] In this system, the origin of the orbital coordinate system is the geocenter O, the ξ-axis coincides with the radius vector of the ascending node N, the ζ-axis coincides with the orbital angular momentum vector, the η-axis is determined by the right-hand rule, and the ξη plane is the orbital plane.

[0041] (2c) Based on Calculate the cosine of the angle β between s and the orbital plane:

[0042]

[0043] in, and They are Components along the ξ and η axes;

[0044] (2d) Calculate the angle between the satellite position vector r and s at the moment of shadow point. Approximate value of cosine:

[0045]

[0046] Where R is the average radius of the Earth;

[0047] (2e) Based on cosβ and Calculate the estimated angle u1 between the projection of s onto the orbital plane and OD using the spherical trigonometric formula, where point D is the exit point of the shadow, and π / 2 < u1 < π.

[0048]

[0049] (2f) Calculate the angle α between the projection of s onto the track plane and ON, where 0 ≤ α < 2π.

[0050]

[0051] (2g) Calculate the estimated time of shadow t based on u1 and α respectively. D_est And the estimated value of the entry time t C_est :

[0052] t D_est =t+(α-u1-u0) / n,

[0053] t C_est =t+(α+u1-u0) / n,

[0054] Among them, satellite orbital angular velocity μ=398600.44km 3 / s 2 It is the gravitational constant.

[0055] Furthermore, the shadow emergence and emergence time determination module iteratively solves for the precise value t of the shadow emergence time using a binary search method. D And the precise value of the projection time t C Specifically, it includes:

[0056] (3a) Based on the relative positions of the Sun, Earth, and satellite at the entrance point C and exit point D, we obtain:

[0057]

[0058] Establish t D and t C The nonlinear solution equation is: g(t) = 0;

[0059] in, t represents any moment in the satellite's orbital flight, r(t) is the position vector of the satellite in the geocentric inertial coordinate system at moment t obtained by analytical extrapolation, and r(t) is the length of r(t);

[0060] (3b) Set t D The solution interval is [t] D_est -p,t D_est +p],t C The solution interval is [t] C_est -p,t C_est +p], the upper and lower bounds of the two intervals are uniformly represented by [t]. low , t up ]express;

[0061] (3c) Calculate t mid =(t up +t low ) / 2;

[0062] (3d) If t up -t low ≤ε,t mid As an optimization result, the emergence time t D =t mid , the moment of entering the film C =t mid Stop the calculation; otherwise, if g(t) low )g(t mid If t < 0, let t up =t mid If g(t) low )g(t mid If t > 0, let t low =t mid Then return to step (3c) to proceed to the next iteration; where ε represents the level of precision.

[0063] Furthermore, the accuracy level ε = 0.5s.

[0064] Furthermore, for orbits of 500–700 km, the maximum duration T of the satellite shadow region is... Umax Take 2400s.

[0065] The advantages of this invention compared to the prior art are:

[0066] (1) This invention provides a method and apparatus for rapid prediction of shadow times in autonomous mission planning for near-Earth remote sensing satellites. Based on the relative positional relationship between the Sun, Earth, and satellite, an innovative estimation model for shadow entry / exit times is established. For each orbital circle of the satellite, firstly, the estimated values ​​of shadow entry and exit are obtained through the estimation model. Then, search intervals are determined centered on the two estimated values. Finally, the bisection method is used to iteratively solve the nonlinear equations satisfied by the shadow entry / exit times, yielding accurate results for the shadow entry and exit times. This method does not involve complex trigonometric function calculations and effectively compresses the length of the iterative solution interval through advance estimation, significantly reducing the computational load. On existing onboard processors, predicting the shadow time for one day takes approximately 0.2 seconds, which is about 100 times faster than existing onboard prediction algorithms, meeting the high real-time requirements of onboard autonomous mission planning.

[0067] (2) The accuracy of the final iterative solution of the method of the present invention mainly depends on the accuracy of the analytical extrapolation of the second-order perturbation of the satellite orbit. The deviation between the calculation results of the entry and exit points of the shadow and the true value can be kept within 2 seconds in a day. The calculation accuracy is nearly twice that of the existing on-board prediction algorithm, which can meet the requirements of the calculation time accuracy of the on-board autonomous mission planning.

[0068] This invention has been applied in orbit in the autonomous mission planning of multiple near-Earth remote sensing satellites, verifying the feasibility and effectiveness of the method. The engineering technology is easy to implement, thus it has practicality. Attached Figure Description

[0069] Figure 1 A flowchart of the rapid shadow forecasting steps provided for the invention;

[0070] Figure 2 A schematic diagram of a cylindrical shadow model provided for the invention;

[0071] Figure 3 A schematic diagram of the projection time prediction model provided for the invention. Detailed Implementation

[0072] The specific embodiments of the present invention will be described in detail below.

[0073] This invention proposes a method and apparatus for predicting the shadow of a near-Earth remote sensing satellite's autonomous mission planning, which can rapidly predict the time of entering and exiting the shadow.

[0074] Assuming the satellite is a near-Earth orbit remote sensing satellite, and sunlight is considered as parallel light, a cylindrical shadow model is established based on the spatial relationship between the sun, earth, and satellite, as follows: Figure 2As shown. Point O is the Earth's center, point C is the point of entry into the shadow, point D is the point of exit from the shadow, S is the solar direction vector in the geocentric inertial coordinate system, DQ is tangent to the Earth's surface, Q is the point of tangency, r is the satellite position vector in the geocentric inertial coordinate system, and R is the Earth's average radius, which is 6371.004 km.

[0075] A satellite orbit circle is defined as follows: the emergence point D is used as the dividing point between adjacent orbit circles, and each orbit circle starts from the emergence point D of the satellite.

[0076] Given that the planning period for the autonomous mission is [T0, T1], it is necessary to predict the Earth's shadow time for all orbital orbits within this planning period. For example... Figure 1 As shown, the steps for rapid land shadow prediction in autonomous mission planning of near-Earth remote sensing satellites are as follows:

[0077] (1) Initialize t0 to the start time T0 of the task planning cycle;

[0078] (2) Based on the relative positional relationship between the Sun, Earth, and the satellite, estimate the emergence time t of the satellite in the orbital sphere to which the satellite belongs at time t0. D_est and the moment of entry into the film t C_est The estimation process is as follows:

[0079] (2a) Based on the latest instantaneous satellite inertial orbital elements obtained from ground-based data or GPS, calculate the average orbital elements of the satellite, and then extrapolate the instantaneous orbital elements of the satellite at time t0 using analytical methods: semi-major axis a0, right ascension of the intersection Ω0, orbital inclination i0, perigee argument ω0 and true perigee angle f0, latitude argument u0=ω0+f0, and simultaneously calculate the satellite's inertial position vector r0 at time t0;

[0080] (2b) By calculating the orbital elements of the Sun in the inertial frame, the unit vector s of the Sun's direction in the geocentric inertial frame at time t0 can be obtained, and then transformed into the orbital coordinate system O-ξηζ to obtain

[0081]

[0082] like Figure 3 As shown, the origin of the orbital coordinate system is the geocenter O, the ξ axis coincides with the radius vector of the ascending node N, the ζ axis coincides with the orbital angular momentum vector, the η axis is determined by the right-hand rule, and the ξη plane is the orbital plane.

[0083] (2c) Using the results obtained in step (2b) Calculate the cosine of the angle β between s and the orbital plane:

[0084]

[0085] in, and They are Components along the ξ and η axes;

[0086] (2d) Calculate the angle between the satellite position vector r and s at time D when the shadow point is located. Approximate value of cosine:

[0087]

[0088] (2e) Based on the calculation results of steps (2c) and (2d), calculate the estimated angle u1 (π / 2 < u1 < π) between the projection of s on the orbital plane and OD according to the spherical trigonometric formula:

[0089]

[0090] (2f) Calculate the angle α between the projection of s onto the track plane and ON (0≤α<2π):

[0091]

[0092] (2g) Calculate the estimated time of shadow t based on the results of steps (2e) and (2f). D_est And the estimated value of the entry time t C_est :

[0093] t D_est =t+(α-u1-u0) / n,

[0094] t C_est =t+(α+u1-u0) / n,

[0095] Among them, satellite orbital angular velocity μ=398600.44km 3 / s 2 It is the gravitational constant.

[0096] (3) The estimated time of emergence t obtained in step (2) D_est And the estimated value of the entry time t C_est The precise value t of the shadow time is obtained by iteratively solving the bisection method in the vicinity. D The precise value t of the projection time C The solution process is as follows:

[0097] (3a) such as Figure 2 As shown, based on the relative positions of the Sun, Earth, and satellite at points C and D, we can obtain:

[0098]

[0099] Therefore, establish t D and t CThe nonlinear solution equation is: g(t) = 0;

[0100] in, t represents any moment in the satellite's orbital flight, r(t) is the position vector of the satellite in the geocentric inertial coordinate system at moment t obtained by analytical extrapolation, and r(t) is the length of r(t);

[0101] (3b) Set t D The solution interval is [t] D_est -p,t D_est +p],t C The solution interval is [t] C_est -p,t C_est +p], where the duration p is the search range centered on the estimated value, p can be 90s, and the upper and lower bounds of the two intervals are uniformly represented by [t]. low ,t up This indicates that steps (3c) and (3d) are used to iteratively solve for the emergence time t in the corresponding solution interval. D and the moment of entry into the film t C ;

[0102] (3c) Calculate t mid =(t up +t low ) / 2;

[0103] (3d) If t up -t low ≤ε,t mid This is the optimization result, which is the iterative solution to the emergence time t of the current orbit. D and the moment of entry into the film t C , t D =t mid t C =t mid Stop the calculation; otherwise, if g(t) low )g(t mid If t < 0, let t up =t mid If g(t) low )g(t mid If t > 0, let t low =t mid Then return to step (3c) to enter the next iteration; where ε represents the precision level, which can be taken as ε = 0.5s.

[0104] (4) Let t0 = t C +T Umax +q, to ​​ensure that you can jump to the next orbital circle by crossing the shaded area of ​​the current orbital circle, where the duration q can be 100s, and T umaxIt is the maximum duration of the satellite's shadow region, for example, in an orbit of 500–700 km. Umax 2400s is acceptable;

[0105] (5) Since it is necessary to calculate the shadow of all orbits within the planning period [T0, T1], determine whether t0 < T1 is satisfied. If so, it means that the calculation has not been completed yet, and go to step (2) to continue calculating the entry and exit time of the next orbit; otherwise, the entry and exit times of all orbits corresponding to the autonomous mission planning period have been calculated and the process ends.

[0106] The present invention also proposes a near-Earth remote sensing satellite autonomous mission planning shadow prediction device, which includes: an initialization module, an exit and entry time estimation module, an exit and entry time determination module, an initialization time determination module, and a mission planning module.

[0107] The initialization module determines the planning period of the autonomous task planning as [T0, T1], and takes the start time T0 of the task planning period as the initialization time t0.

[0108] The emergence and exit time estimation module calculates the estimated emergence time t of the satellite in its orbital sphere at time t0 based on the relative positions of the Sun, Earth, and the satellite. D_est And the estimated value of the entry time t C_est .

[0109] The exit and entry time determination module is based on the estimated exit time t. D_est And the estimated value of the entry time t C_est The precise values ​​t for the emergence and emergence times are obtained by iteratively solving using the bisection method. D and t C .

[0110] The initialization timing determination module determines the precise value t of the emergence timing. C Maximum duration T of satellite shadow area umax The sum of these values ​​plus the duration q is used as the final initialization time t0, where the duration q can be taken as 100s. For orbits of 500–700 km, the maximum duration T of the satellite shadow region is... Umax Take 2400s.

[0111] The task planning module determines whether the final initialization time t0 is less than the termination time T1 of the task planning cycle. If so, the exit and entry time estimation module, the exit and entry time determination module, and the initialization time determination module continue to calculate the exit and entry time of the next orbital circle; otherwise, the exit and entry times of all orbital circles corresponding to the autonomous task planning cycle have been calculated.

[0112] The emergence and emergence time estimation module calculates the predicted emergence time t of the satellites in the orbital sphere to which the satellite belongs at time t0. D_est And the predicted time of advance tC_est Specifically, it includes:

[0113] (2a) Determine the instantaneous orbital elements of the satellite at time t0: semi-major axis a0, right ascension of the intersection Ω0, orbital inclination i0, perigee argument ω0 and true perigee argument f0, latitude argument u0=ω0+f0, and calculate the satellite's inertial position vector r0 at time t0.

[0114] (2b) The unit vector s of the sun's direction in the geocentric inertial frame at time t0 is calculated by calculating the orbital elements of the sun in the inertial frame, and then transformed into the orbital coordinate system O-ξηζ to obtain the following result.

[0115]

[0116] In this system, the origin of the orbital coordinate system is the Earth's center O, the ξ-axis coincides with the radius vector of the ascending node N, the ζ-axis coincides with the orbital angular momentum vector, the η-axis is determined by the right-hand rule, and the ξη plane is the orbital surface.

[0117] (2c) Based on Calculate the cosine of the angle β between s and the orbital plane:

[0118]

[0119] in, and They are Components along the ξ and η axes.

[0120] (2d) Calculate the angle between the satellite position vector r and s at the moment of shadow point. Approximate value of cosine:

[0121]

[0122] Where R is the average radius of the Earth.

[0123] (2e) Based on cosβ and Calculate the estimated angle u1 between the projection of s onto the orbital plane and OD using the spherical trigonometric formula, where point D is the exit point of the shadow, and π / 2 < u1 < π.

[0124]

[0125] (2f) Calculate the angle α between the projection of s onto the track plane and ON, where 0 ≤ α < 2π:

[0126]

[0127] (2g) Calculate the estimated time of shadow t based on u1 and α respectively. D_est And the estimated value of the entry time t C_est :

[0128] t D_est =t+(α-u1-u0) / n,

[0129] t C_est =t+(α+u1-u0) / n,

[0130] Among them, satellite orbital angular velocity μ=398600.44km 3 / s 2 It is the gravitational constant.

[0131] The module for determining the emergence and exit times uses a bisection method to iteratively solve for the precise value t of the emergence time. D The precise value t of the projection time C Specifically, it includes:

[0132] (3a) Based on the relative positions of the Sun, Earth, and satellite at the entrance point C and exit point D, we obtain:

[0133]

[0134] Establish t D and t C The nonlinear solution equation is: g(t)=0.

[0135] in, t represents any moment in the satellite's orbital flight, r(t) is the position vector of the satellite in the geocentric inertial coordinate system at moment t obtained by analytical extrapolation, and r(t) is the length of r(t).

[0136] (3b) Set t D The solution interval is [t] D_est -p,t D_est +p],t C The solution interval is [t] C_est -p,t C_est +p], where the duration p can be 90s, and the upper and lower bounds of the two intervals are uniformly represented by [t]. low , t up ]express.

[0137] (3c) Calculate t mid =(t up +t low ) / 2.

[0138] (3d) If t up -t low ≤ε,t mid As an optimization result, the emergence time t D =t mid , the moment of entering the film C =tmid Stop the calculation; otherwise, if g(t) low )g(t mid If t < 0, let t up =t mid If g(t) low )g(t mid If t > 0, let t low =t mid Then return to step (3c) to proceed to the next iteration; where ε represents the precision level, and ε can be set to 0.5s.

[0139] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for land shadow prediction in autonomous mission planning of near-Earth remote sensing satellites, characterized in that, include: (1) Determine the planning period of the autonomous task planning as [T0, T1], and take the start time T0 of the task planning period as the initialization time. t 0; (2) Based on the relative positions of the Sun, Earth, and satellite, calculate t Estimated emergence time t of the satellite in the orbital sphere at time 0 D_est And the estimated value of the entry time t C_est ; (3) Based on the estimated time of emergence t D_est and estimated time of entry t C_est The precise value of the shadow time was obtained by iteratively solving using the bisection method. t D Precise value of the time of entry t C ; In step (3), the precise value of the shadow time is obtained by iteratively solving the problem using the bisection method. t D and the precise value of the moment of entry. t C Specifically, it includes: (3a) Based on the relative positions of the entrance point C, the exit point D, and the Sun-Earth-Satellite, we obtain: Establish t D and t C Nonlinear solution equations: = 0; in, t This represents any moment in the satellite's orbit. For parsing extrapolation t The position vector of the satellite in the geocentric inertial coordinate system at any given time. It is a position vector Length; R It is the Earth's average radius; for t The unit vector of the sun's direction in the geocentric inertial frame at time 0; The satellite position vector at the time of emergence. and The included angle; (3b) Settings t D The solution interval is [ t D_est - p , t D_est + p ], t C The solution interval is [ t C_est - p , t C_est + p The upper and lower bounds of the two intervals are uniformly represented by []. t low , t up [Indicates duration] p The search range is centered on the estimated value; (3c) Calculation ; (3d) If , As an optimization result, the time of appearance Moment of filming Stop the calculation; otherwise, if ,make ,like ,make Then return to step (3c) to proceed to the next iteration; where, Indicates the level of precision; (4) The precise value of the advance time t C Maximum duration T of satellite shadow area Umax The sum plus time margin q As the final initialization moment t 0; (5) Determine the final initialization time t Is 0 less than the end time of the task planning cycle? If so, proceed to step (2) to continue calculating the entry and exit times of the next orbital circle; otherwise, the entry and exit times of all orbital circles corresponding to the autonomous mission planning cycle have been calculated.

2. The method according to claim 1, characterized in that, The calculation described in step (2) t Estimated emergence time t of the satellite in the orbital sphere at time 0 D_est And the estimated value of the entry time t C_est Specifically, it includes: (2a) Determine t The instantaneous orbital elements of the satellite at time 0, including the semi-major axis. Right ascension of the intersection Track inclination Perigeal argument And true near point angle Latitude angle and calculate t Satellite inertial position vector at time 0 r 0; (2b) Calculate by calculating the orbital elements of the Sun in the inertial frame. t The unit vector of the sun's direction in the geocentric inertial frame at time 0. and transform it to the orbital coordinate system. get : ； In this system, the origin O of the orbital coordinate system is the Earth's center. The axis coincides with the radius vector of the ascending node N. The axis coincides with the orbital angular momentum vector. The axis is determined by the right-hand rule. The plane is the track surface; (2c) Based on calculate Angle with the track surface cosine value: ， in, and They are exist shaft and The components of the axis; (2d) Calculate the satellite position vector at the time of shadow point. and included angle Approximate value of cosine: ； (2e) Based on and Calculate based on spherical trigonometric formulas Estimated angle between the projection onto the orbital plane and the OD. Point D is the point where the shadow appears. , ； (2f) Calculation The angle between the projection onto the track plane and ON , , ； (2g) based and Calculate the estimated time of shadow respectively t D_est and estimated time of entry t C_est : ， ， Among them, satellite orbital angular velocity , It is the gravitational constant.

3. The method according to claim 1, characterized in that, The level of accuracy .

4. The method according to claim 1, characterized in that, For orbits of 500–700 km, the maximum duration T of the satellite shadow region is... Umax Take 2400s.

5. A near-Earth remote sensing satellite autonomous mission planning and ground shadow prediction device, characterized in that, include: The module includes an initialization module, an entry / exit time estimation module, an entry / exit time determination module, and an initialization time determination module. Among these, The initialization module determines the planning period of the autonomous task planning as [T0, T1], and uses the start time T0 of the task planning period as the initialization time. t 0; The solar phase estimating module calculates the time of solar phase emergence based on the relative positions of the Sun, Earth, and satellite. t Estimated emergence time t of the satellite in the orbital sphere at time 0 D_est And the estimated value of the entry time t C_est ; The exit and entry time determination module is based on the estimated exit time value. t D_est and estimated time of entry t C_est The precise value of the shadow time was obtained by iteratively solving using the bisection method. t D Precise value of the time of entry t C ; The initialization timing determination module determines the precise value of the projection timing. t C Maximum duration T of satellite shadow area Umax The sum plus time margin q As the final initialization moment t 0; The task planning module determines the final initialization time. t Is 0 less than the end time of the task planning cycle? If so, the exit and entry time estimation module, the exit and entry time determination module, and the initialization time determination module continue to calculate the exit and entry time of the next orbital cycle; otherwise, the exit and entry times of all orbital cycles corresponding to the autonomous mission planning cycle have been calculated. The shadow emergence and emergence time determination module uses a binary search method to iteratively solve for the precise value of the shadow emergence time. t D Precise value of the time of entry t C Specifically, it includes: (3a) Based on the relative positions of the entrance point C, the exit point D, and the Sun-Earth-Satellite, we obtain: Establish t D and t C Nonlinear solution equations: = 0; in, t This represents any moment in the satellite's orbit. For parsing extrapolation t The position vector of the satellite in the geocentric inertial coordinate system at any given time. It is a position vector Length; R It is the Earth's average radius; for t The unit vector of the sun's direction in the geocentric inertial frame at time 0; The satellite position vector at the time of emergence. and The included angle; (3b) Settings t D The solution interval is [ t D_est - p , t D_est + p ], t C The solution interval is [ t C_est - p , t C_est + p The upper and lower bounds of the two intervals are uniformly represented by []. t low , t up [Indicates duration] p The search range is centered on the estimated value; (3c) Calculation ; (3d) If , As an optimization result, the time of appearance Moment of filming Stop the calculation; otherwise, if ,make ,like ,make Then return to step (3c) to proceed to the next iteration; where, Indicates the level of precision.

6. The apparatus according to claim 5, characterized in that, The exit and entrance time estimation module calculates... t The predicted emergence time t of the satellite in the orbital sphere to which time 0 belongs D_est And the predicted time of advance t C_est Specifically, it includes: (2a) Determine t The instantaneous orbital elements of the satellite at time 0, including the semi-major axis. Right ascension of the intersection Track inclination Perigeal argument And true near point angle Latitude angle and calculate t Satellite inertial position vector at time 0 r 0; (2b) Calculate by calculating the orbital elements of the Sun in the inertial frame. t The unit vector of the sun's direction in the geocentric inertial frame at time 0. and transform it to the orbital coordinate system. get : ； The origin of the orbital coordinate system is the Earth's center O. The axis coincides with the radius vector of the ascending node N. The axis coincides with the orbital angular momentum vector. The axis is determined by the right-hand rule. The plane is the track surface; (2c) Based on calculate Angle with the track surface cosine value: ， in, and They are exist shaft and The components of the axis; (2d) Calculate the satellite position vector at the time of shadow point. and included angle Approximate value of cosine: ； (2e) Based on and Calculate based on spherical trigonometric formulas Estimated angle between the projection onto the orbital plane and the OD. Point D is the point where the shadow appears. , ； (2f) Calculation The angle between the projection onto the track plane and ON , , ； (2g) based and Calculate the estimated time of shadow respectively t D_est and estimated time of entry t C_est : ， ， Among them, satellite orbital angular velocity , It is the gravitational constant.

7. The apparatus according to claim 5, characterized in that, The level of accuracy .

8. The apparatus according to claim 5, characterized in that, For orbits of 500–700 km, the maximum duration T of the satellite shadow region is... Umax Take 2400s.

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