A method, apparatus and medium for launching satellites in emergency scenarios

Abstract

This invention provides a method, apparatus, and medium for launching satellites in emergency scenarios, comprising: if the satellite launch point is a fixed value and the orbital inclination is an unknown value, determining the first latitude argument and the geographical longitude of the first ascending node when the satellite passes through the launch point at the first current orbital inclination, and the second latitude argument and the geographical longitude of the second ascending node when passing through the observation point, based on the coordinates of the observation point and the coordinates of the launch point; determining the target orbital inclination of the satellite and the corresponding launch window based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node; if the satellite launch point is an unknown value and the orbital inclination is a fixed value, determining the longitude and latitude of the nadir point corresponding to each launch window based on the coordinates of the observation point, the target observation time, and the observation time step; thus, in emergency launch scenarios, the launch window can be determined according to the launch requirements, thereby quickly and accurately launching the satellite to a preset position to achieve fixed-point observation of emergency events.

Description

Technical Field

[0001] This invention relates to the field of satellite launch technology, and in particular to a method, apparatus and medium for launching satellites based on emergency scenarios. Background Technology

[0002] When encountering major emergencies such as natural disasters, hotspot conflicts, and regional information enhancement, launching spacecraft into space in an emergency to quickly conduct fixed-point observations of designated locations is an important means of disaster reconnaissance, hotspot monitoring, and command and communication.

[0003] Therefore, how to accurately launch satellites into space so as to quickly achieve fixed-point observation of designated locations is a technical problem that needs to be solved. Summary of the Invention

[0004] To address the problems existing in the prior art, embodiments of the present invention provide a method, apparatus, and medium for launching satellites based on emergency scenarios, in order to solve or partially solve the technical problem in the prior art that it is impossible to quickly and accurately launch satellites to a predetermined location in space when a fixed-point observation is required during an emergency event.

[0005] A first aspect of the present invention provides a method for launching a satellite based on an emergency scenario, the method comprising:

[0006] If the launch point of the satellite is determined to be a fixed value and the orbital inclination is an unknown value, the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point under the first current orbital inclination are determined according to the coordinates of the observation point and the coordinates of the launch point. The second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point under the first current orbital inclination are determined according to the coordinates of the observation point and the coordinates of the launch point.

[0007] The target orbital inclination of the satellite and the corresponding launch window are determined based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node.

[0008] If the launch point of the satellite is determined to be an unknown value and the orbital inclination is a fixed value, the longitude and latitude of the nadir point under each launch window are determined according to the coordinates of the observation point, the target observation time, and the preset observation time step.

[0009] In the above scheme, the first latitude argument includes: a first sub-latitude argument and a second sub-latitude argument; the first ascending node longitude includes: a first sub-ascending node longitude and a second sub-ascending node longitude; determining the first latitude argument and the first ascending node longitude corresponding to the launch point at the first current orbital inclination, based on the observation point coordinates and the launch point coordinates, includes:

[0010] If it is determined that the satellite's orbital descent passes through the launch point, according to the formula... Determine the first sub-latitude argument u0 corresponding to the satellite's descent orbit passing the launch point at the first current orbital inclination angle;

[0011] The geographical longitude Ω0 of the first sub-ascending node corresponding to the satellite when the first current orbital inclination descends through the launch point is determined according to the formula Ω0=L0-arctan(cos(i)·tan(u0))-π.

[0012] If it is determined that the satellite's orbital ascent passes through the launch point, according to the formula... Determine the second sub-latitude argument u′0 corresponding to the satellite's ascent and passage through the launch point at the first current orbital inclination;

[0013] The geographical longitude Ω′0 of the second sub-ascending node corresponding to the satellite's ascent at the launch point under the first current orbital inclination is determined according to the formula Ω′0=L0-arctan(cos(i)·tan(u0)); where Ω′0=L0-arctan(cos(i)·tan(u0))

[0014] B0 is the latitude of the launch point, i is the first current orbital inclination, i∈[B_m,π-B_m], B_m=max(|B0|,|B1|), B1 is the latitude of the observation point, and L0 is the longitude of the launch point.

[0015] In the above scheme, the second latitude argument includes: a third sub-latitude argument and a fourth sub-latitude argument; the second ascending node longitude includes: a third sub-ascending node longitude and a fourth sub-ascending node longitude; determining the second latitude argument and the second ascending node longitude corresponding to the satellite passing the observation point at the first current orbital inclination includes:

[0016] If it is determined that the satellite's orbital descent passes through the observation point, according to the formula... Determine the third sub-latitude argument u1 corresponding to the observation point when the satellite's descending orbit at the first current orbital inclination angle passes through the observation point;

[0017] The geographical longitude Ω1 of the third sub-ascending node corresponding to the observation point when the satellite's descending orbit at the first current orbital inclination passes through the observation point is determined according to the formula Ω1=L1-arctan(cos(i)·tan(u1))-π.

[0018] If it is determined that the satellite's orbital ascent passes through the observation point, according to the formula... Determine the fourth sub-latitude argument u1′ corresponding to the observation point when the satellite ascends to its orbit at the first current orbital inclination.

[0019] The geographical longitude Ω′1 of the fourth sub-ascending node corresponding to the observation point when the satellite ascends to its first current orbital inclination is determined according to the formula Ω′1=L1-arctan(cos(i)·tan(u1)); where Ω′1=L1-arctan(cos(i)·tan(u1))

[0020] Where i is the first current orbital inclination angle, and i ∈ [B_m, π-B_m], the... B1 is the latitude of the observation point, B0 is the latitude of the launch point, and L1 is the longitude of the observation point.

[0021] In the above scheme, determining the target orbital inclination of the satellite and the corresponding launch window based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node includes:

[0022] For any combination pattern among all combination patterns, perform the following processing:

[0023] Determine the precession rate W of the right ascension of the satellite's ascending node at the first current orbital inclination. Ω Perimeter angle precession rate W ω and the precession rate W at the approximate point angle M According to the formula Determine the first time t required for the satellite to travel from the first latitude angle to the second latitude angle. u According to the formula The second time t required for the satellite to precess from the first ascending node's geographic longitude to the second ascending node's geographic longitude is determined. Ω If the absolute value of the duration difference between the first duration and the second duration is determined to be less than a preset duration threshold, then the first current orbital inclination angle is determined as the reference orbital inclination angle in the combined mode; and T0-t u The reference launch window is determined as the reference orbit inclination angle in the combined mode; the combined mode includes: the mode in which the satellite descends to a lower orbit and passes through the launch point and the observation point, the mode in which the satellite ascends to a higher orbit and passes through the launch point and the observation point, the mode in which the satellite ascends to a higher orbit and passes through the launch point and the satellite descends to a lower orbit and the observation point, and the mode in which the satellite descends to a lower orbit and passes through the launch point and the satellite ascends to a higher orbit and passes through the observation point;

[0024] The minimum value among all reference orbit inclinations obtained under all combination modes is determined as the target orbit inclination of the satellite, and the reference launch window corresponding to the target orbit inclination is the target launch window; wherein...

[0025] Δu is the difference in latitude argument between the first latitude argument and the second latitude argument, and ΔΩ is the difference in geographical longitude between the first ascending node and the second ascending node.

[0026] In the above scheme, if the absolute value of the duration difference between the first duration and the second duration is not less than a preset duration threshold, then the orbital inclination angle is increased according to the orbital inclination angle step size. Each time it is increased, the following processing needs to be performed:

[0027] Under the increased orbital inclination, determine the first time required for the satellite to travel from the first latitude argument to the second latitude argument, and the second time required for the satellite to travel from the first ascending node longitude to the second ascending node longitude; determine the time difference between the first and second times under the increased orbital inclination; further determine whether the absolute value of the time difference under the increased orbital inclination is less than the time threshold. If it is less, then the increased orbital inclination is determined as the reference orbital inclination in the combined mode; if it is not less, then determine whether the time difference under the increased orbital inclination has the same sign as the time difference under the original orbital inclination.

[0028] If the signs are the same, continue to increase the track inclination angle and repeat the above steps; repeat this process until the increased track inclination angle i is determined. m The difference in duration and the increase in the orbital inclination angle i n If the time difference values ​​are of opposite signs, then the inclination angle of the reference orbit is determined to be within [i]. n i m Within the interval, according to the increased track inclination angle i m and the increased front track inclination angle i n Determine the inclination angle of the reference orbit.

[0029] In the above scheme, based on the increased track inclination angle i m and the increased front track inclination angle i n Determining the inclination angle of the reference orbit includes:

[0030] Step S1: Set the first track inclination angle i1 and the second track inclination angle i2; the initial value of the first track inclination angle i1 is the track inclination angle i2 before the increase. n The initial value of the second track inclination angle i2 is the increased track inclination angle i. m ;

[0031] Step S2: Determine the second current orbital inclination i′ as the median of the first orbital inclination and the second orbital inclination, i′ = (i1 + i2) / 2; Under the second current orbital inclination, determine the first time required for the satellite to travel from the first latitude argument to the second latitude argument and the second time required for the satellite to travel from the geographical longitude of the first ascending node to the geographical longitude of the second ascending node; Determine the time difference between the first time and the second time under the second current orbital inclination; Determine whether the absolute value of the time difference between the first time and the second time under the second current orbital inclination is less than a preset time threshold. If it is less, then determine the second current orbital inclination as the reference orbital inclination.

[0032] Step S3: If the absolute value of the duration difference between the first duration and the second duration under the second current track inclination is not less than the duration threshold, then determine the track inclination angle that corresponds to the duration difference with the opposite sign under the second current track inclination angle i1 and the second track inclination angle i2; reassign the two track inclination angles to the first track inclination angle i1 and the second track inclination angle i2; reassign the two track inclination angles with the opposite sign of the duration difference to the first track inclination angle i1 and the second track inclination angle i2.

[0033] Repeat steps S2 and S3 until the absolute value of the time difference corresponding to the second current orbital inclination is less than the preset time threshold. Then, determine the second current orbital inclination whose absolute value of the time difference is less than the time threshold as the reference orbital inclination in the combined mode.

[0034] In the above scheme, determining the longitude and latitude of the nadir point corresponding to each launch window based on the observation point coordinates, the target observation time, and the preset observation time step includes:

[0035] For any combination pattern among all combination patterns, perform the following processing:

[0036] Obtain the predetermined orbital inclination of the satellite, and determine the third latitude argument and the geographical longitude of the third ascending node when the satellite passes through the observation point based on the predetermined orbital inclination;

[0037] According to the formula Ω t =Ω2+(W′) Ω -ω e The geographic longitude Ω of the current ascending node is determined by )×nΔ′t when the satellite moves back n observation time steps Δ′t from the observation time of the target. t ;

[0038] According to formula u t =u2+(W′) ω +W′ MMThe current latitude argument u is determined by )×nΔ′t when the satellite moves back n observation time steps Δ′t from the observation time of the target. t ;

[0039] Determine the geographic latitude of the nadir point when the satellite regresses n observation time steps from the target observation time based on the current latitude argument;

[0040] Based on the current ascending node's geographic longitude, determine the nadir longitude of the satellite when it regresses n observation time steps from the target observation time; where...

[0041] Ω2 is the third latitude argument, u2 is the geographical longitude of the third ascending node, and W′ is... Ω ω is the precession rate of the right ascension of the ascending node of the satellite at the preset orbital inclination. e The angular velocity of Earth's rotation, W′ ω W′ is the perigee argument precession rate of the satellite at the preset orbital inclination. M The precession rate of the satellite at the mean apogee angle under the preset orbital inclination is given.

[0042] In the above scheme, determining the geographic latitude of the nadir point when the satellite regresses n observation time steps from the target observation time based on the current latitude argument includes:

[0043] According to formula B t =arcsin(sin(u t The geographic latitude B of the nadir point is determined by sin(i″)) when the satellite moves back n observation time steps from the observation time of the target. t ;in,

[0044] The u t The current latitude argument is the angle at which the satellite moves backward n time steps from the observation time of the target, and i″ is the preset orbital inclination angle.

[0045] In the above scheme, determining the nadir longitude of the satellite when it has regressed n observation time steps from the target observation time based on the current ascending node longitude includes:

[0046] According to the formula Determine the geographic longitude L of the nadir point when the satellite regresses n observation time steps from the target observation time. t ;in,

[0047] The Ω t The geographical longitude of the current ascending node of the satellite is defined as n observation time steps Δ′t, which is the time step backward from the observation time of the target. The preset orbital inclination is defined as u. tThe current latitude argument is the angle at which the satellite moves back n observation time steps Δ′t from the observation time of the target.

[0048] A second aspect of the present invention provides an apparatus for launching satellites based on emergency scenarios, the apparatus comprising:

[0049] The first determining unit is configured to, if the launch point of the satellite is determined to be a fixed value and the orbital inclination is an unknown value, determine the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point under the first current orbital inclination, based on the coordinates of the observation point and the coordinates of the launch point; and determine the second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point under the first current orbital inclination, based on the coordinates of the observation point and the coordinates of the launch point.

[0050] The second determining unit is used to determine the target orbital inclination of the satellite and the corresponding launch window based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node;

[0051] The third determining unit is used to determine the longitude and latitude of the nadir point corresponding to each launch window if the launch point of the satellite is an unknown value and the orbital inclination is a fixed value, based on the coordinates of the observation point, the target observation time, and the preset observation time step.

[0052] A third aspect of the present invention provides a computer-readable storage medium having a computer program stored thereon that, when executed by a processor, implements the steps of the method described in any of the first aspects.

[0053] This invention provides a method, apparatus, and medium for launching a satellite based on an emergency scenario. The method includes: if the launch point of the satellite is determined to be a fixed value and the orbital inclination is an unknown value, determining, based on the coordinates of the observation point and the coordinates of the launch point, the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point at a first current orbital inclination; and determining, based on the coordinates of the observation point and the coordinates of the launch point, the second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point at the first current orbital inclination; and determining, based on the first latitude argument, the geographical longitude of the first ascending node, and the coordinates of the launch point, the second latitude argument and the second ascending node geographical longitude corresponding to the satellite passing through the observation point at the first current orbital inclination; and determining, based on the first latitude argument, the first ascending node geographical longitude, and the second ascending node geographical longitude, the satellite launch point is launched at an emergency scenario. The target orbital inclination of the satellite and the corresponding launch window are determined by the second latitude argument and the geographic longitude of the second ascending node. If the launch point of the satellite is unknown and the orbital inclination is fixed, the longitude and latitude of the nadir point under each launch window are determined according to the coordinates of the observation point, the target observation time, and the preset observation time step. In this way, in emergency launch scenarios, the launch window can be determined according to the launch requirements (fixed orbital inclination or fixed launch point), and the satellite can be launched quickly and accurately to the preset position based on the launch window and launch point to achieve fixed-point observation of emergency events. Attached Figure Description

[0054] Various other advantages and benefits will become apparent to those skilled in the art upon reading the following detailed description of preferred embodiments. The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. Furthermore, the same reference numerals denote the same parts throughout the drawings. In the drawings:

[0055] Figure 1 A schematic flowchart of a method for launching a satellite based on an emergency scenario according to an embodiment of the present invention is shown;

[0056] Figure 2 A schematic diagram of a device for launching satellites based on an emergency scenario according to an embodiment of the present invention is shown. Detailed Implementation

[0057] Exemplary embodiments of the present disclosure will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be implemented in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art.

[0058] This invention provides a method for launching satellites based on emergency scenarios, such as... Figure 1 As shown, the method mainly includes the following steps:

[0059] S110, if the launch point of the satellite is determined to be a fixed value and the orbital inclination is an unknown value, the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point under the first current orbital inclination are determined according to the coordinates of the observation point and the coordinates of the launch point; and the second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point under the first current orbital inclination are determined according to the coordinates of the observation point and the coordinates of the launch point.

[0060] There are two launch modes when launching satellites. The first is that the launch point is fixed but the orbital inclination is not fixed. The second is that the launch point is movable (not fixed) but the orbital inclination is fixed.

[0061] When an emergency occurs, either the first or second launch mode can be selected for rapid satellite launch within a limited time. The methods for determining the launch window differ between the first and second launch modes. Therefore, in this invention, if the satellite launch point is a fixed value and the orbital inclination is unknown, the first latitudinal argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point at the first current orbital inclination are determined based on the coordinates of the observation point and the launch point. Furthermore, the second latitudinal argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point at the first current orbital inclination are determined based on the coordinates of the observation point and the launch point.

[0062] Specifically, a satellite can either descend into a lower orbit or ascend into a higher orbit past the launch point. Similarly, a satellite can either descend into a lower orbit or ascend into a higher orbit past the observation point. Therefore, for the launch point, the first latitude argument includes: the first sub-latitude argument (when the satellite descends into a lower orbit past the launch point) and the second sub-latitude argument (when the satellite ascends into a higher orbit past the launch point); the first ascending node's geographical longitude includes: the first sub-ascending node's geographical longitude (when the satellite descends into a lower orbit past the launch point) and the second sub-ascending node's geographical longitude (when the satellite ascends into a higher orbit past the launch point).

[0063] For the observation point, the second latitude argument includes: the third sub-latitude argument (when the satellite descends through the observation point) and the fourth sub-latitude argument (when the satellite ascends through the observation point); the second ascending node longitude includes: the third sub-ascending node longitude (when the satellite descends through the observation point) and the fourth sub-ascending node longitude (when the satellite ascends through the observation point).

[0064] Specifically, the satellite's orbital ascent marker as it passes the launch point can be set to flag0. When flag0 = 0, it indicates that the satellite passes the launch point in a descending orbit; when flag0 = 1, it indicates that the satellite passes the launch point in an ascending orbit. Similarly, the satellite's orbital ascent marker as it passes the observation point can be set to flag1. When flag1 = 0, it indicates that the satellite passes the observation point in a descending orbit; when flag1 = 1, it indicates that the satellite passes the observation point in an ascending orbit.

[0065] In one implementation, determining the first latitude argument and the geographical longitude of the first ascending node corresponding to the launch point at the first current orbital inclination, based on the coordinates of the observation point and the launch point, includes:

[0066] If it is determined that the satellite descends to pass the launch point (flag0 = 0), the first sub-latitude argument u0 corresponding to the satellite descending to pass the launch point at the first current orbital inclination is determined according to formula (1):

[0067]

[0068] According to formula (2), determine the geographical longitude Ω0 of the first sub-ascending node corresponding to the satellite's descent orbit passing through the launch point at the first current orbital inclination angle:

[0069] Ω0=L0-arctan(cos(i)·tan(u0))-π(2)

[0070] If it is determined that the satellite ascends to pass through the launch point (flag0 = 1), the second sub-latitude argument u′0 corresponding to the satellite ascending to pass through the launch point at the first current orbital inclination is determined according to formula (3):

[0071]

[0072] According to formula (4), determine the geographical longitude Ω′0 of the second sub-ascending node when the satellite passes through the launch point at the first current orbital inclination.

[0073] Ω′0=L0-arctan(cos(i)·tan(u0)) (4)

[0074] Where B0 is the latitude of the launch point, i is the first current orbital inclination, i∈[B_m,π-B_m], B_m=max(|B0|,|B1|), B1 is the latitude of the observation point, and L0 is the longitude of the launch point.

[0075] Similarly, in one implementation, determining the second latitudinal argument and the geographical longitude of the second ascending node when the satellite passes the observation point at the first current orbital inclination, based on the coordinates of the observation point and the launch point, includes:

[0076] If it is determined that the satellite descends to pass the observation point (flag1 = 0), the third sub-latitude argument u1 corresponding to the satellite descending to pass the observation point at the first current orbital inclination is determined according to formula (5):

[0077]

[0078] The geographical longitude Ω1 of the third ascending node corresponding to the satellite's descending orbit at the first current orbital inclination when it passes the observation point is determined using the formula:

[0079] Ω1=L1-arctan(cos(i)·tan(u1))-π (6)

[0080] If it is determined that the satellite ascends through the observation point (flag1 = 1), the fourth sub-latitude argument u′1 corresponding to the satellite ascending through the observation point at the first current orbital inclination is determined according to formula (7):

[0081]

[0082] According to formula (8), determine the geographical longitude Ω′1 of the fourth sub-ascending node corresponding to the observation point when the satellite ascends to its first current orbital inclination.

[0083] Ω′1=L1-arctan(cos(i)·tan(u1)) (8)

[0084] Where i is the first current orbital inclination angle, i∈[B_m,π-B_m], B_m=max(|B0|,|B1|), B1 is the latitude of the observation point, B0 is the latitude of the launch point, and L1 is the longitude of the observation point.

[0085] In this way, the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point at the first current orbital inclination can be determined, as well as the second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point.

[0086] S111, determine the target orbital inclination of the satellite and the corresponding launch window based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node.

[0087] Once the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite's passage through the launch point, and the second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite's passage through the observation point are determined, the target orbital inclination of the satellite and the corresponding launch window are determined based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node. The specific implementation is as follows:

[0088] Determine the precession rate W of the right ascension of the ascending node of the satellite at the first current orbital inclination. Ω Perimeter angle precession rate W ω and the precession rate W at the approximate point angle M ;

[0089] The combined modes include: a mode where the satellite descends to its lower orbit and passes through the launch point, a mode where the satellite ascends to its upper orbit and passes through the launch point, a mode where the satellite ascends to its upper orbit and passes through the launch point, and a mode where the satellite descends to its lower orbit and passes through the launch point, and a mode where the satellite descends to its lower orbit and passes through the launch point, and a mode where the satellite ascends to its upper orbit and passes through the observation point. For any combination mode among all these combined modes, the following processing is performed:

[0090] The first time t required for the satellite to travel from the first latitude angle to the second latitude angle is determined according to formula (9). u :

[0091]

[0092] The second time t required for the satellite to precess from the geographical longitude of the first ascending node to the geographical longitude of the second ascending node is determined according to formula (10). Ω :

[0093]

[0094] If the absolute value of the duration difference between the first duration and the second duration is less than a preset duration threshold, then the first current orbital inclination angle is determined as the reference orbital inclination angle in this combination mode; and T0-t u The reference launch window is determined as the reference orbit inclination angle in the combined mode.

[0095] After performing the above processing on all combination modes, the minimum value among the reference orbit inclinations obtained from all combination modes is determined as the target orbit inclination of the satellite, and the reference launch window corresponding to the target orbit inclination is the target launch window; where...

[0096] Δu is the difference in latitudinal argument between the first and second latitude arguments, and ΔΩ is the difference in geographical longitude between the first and second ascending nodes.

[0097] Specifically, in one implementation, determining the satellite's ascending node right ascension precession rate, perigee argument precession rate, and mean perigee precession rate at the first current orbital inclination includes:

[0098] The precession rate W of the right ascension of the ascending node at the first current orbital inclination is determined according to formula (11). Ω :

[0099]

[0100] The perigee angle precession rate W of the satellite at the first current orbital inclination angle is determined according to the formula. ω :

[0101]

[0102] The precession rate W of the satellite at its first current orbital inclination angle is determined using the formula. M :

[0103]

[0104] Where J2 is the coefficient of the J2 term of Earth's gravity, J2 = 0.001082636; A e Let A be the semi-major axis of the Earth ellipsoid model. e =6378137; μ is the Earth's gravitational constant, μ = 3.986004418 × 10 14 , where a is the semi-major axis of the orbit and i is the first current orbital inclination angle.

[0105] For the first current orbital inclination under any combination mode, the first time t required for the satellite to travel from the first latitude argument to the second latitude argument is determined according to formula (14). u :

[0106]

[0107] The second time t required for the satellite to precess from the geographical longitude of the first ascending node to the geographical longitude of the second ascending node is determined according to formula (15). Ω ;

[0108]

[0109] Finally, the duration difference Δt between the first and second durations is determined according to formula (16):

[0110] Δt=t u -t Ω (16)

[0111] Then, it is determined whether |Δt| is less than a preset duration threshold, which can be 0.1s. If it is determined that |Δt| is less than the preset duration threshold, then the launch window corresponding to the first current orbital inclination angle in this combination mode is determined to be T. w =T0-t u T0 represents the time when the satellite observes the observation point.

[0112] For example, suppose the iteration reaches its 8th iteration, and the obtained |Δt| is 0.05. The first current orbital inclination angle corresponding to the 8th calculation is 30°. Since |Δt| is less than 0.1s, the condition is met, and the first duration t at this time will be obtained. u According to formula T w =T0-t u Determine the corresponding launch window.

[0113] Furthermore, if it is determined that the absolute value of the duration difference between the first duration and the second duration is not less than a preset duration threshold, then the orbital inclination angle is increased according to the orbital inclination angle step size. Each time it is increased, the following processing needs to be performed:

[0114] Under the increased orbital inclination, determine the first time required for the satellite to travel from the first latitude argument to the second latitude argument, and the second time required for the satellite to travel from the geographical longitude of the first ascending node to the geographical longitude of the second ascending node; determine the time difference between the first and second times under the increased orbital inclination; further determine whether the absolute value of the time difference under the increased orbital inclination is less than the time threshold. If it is less, then the increased orbital inclination is determined as the reference orbital inclination in the combined mode; if it is not less, then determine whether the time difference under the increased orbital inclination has the same sign as the time difference under the orbital inclination before the increase.

[0115] If the signs are the same, continue to increase the track inclination angle and repeat the above steps; repeat this process until the increased track inclination angle i is determined. m The difference in duration and the increase in the orbital inclination angle i n If the duration difference values ​​are of opposite signs, then the inclination of the reference orbit is determined to be within [i]. n i m Within the interval, based on the increased track inclination angle i m and increase the forward inclination angle i n Determine the inclination angle of the reference orbit.

[0116] In this context, "same sign" means that the time difference under the increased track inclination angle and the time difference under the original track inclination angle are both positive or both negative. "Odd signs" means that one of the time difference under the increased track inclination angle and the other under the original track inclination angle is positive and the other is negative.

[0117] In one alternative implementation, based on the increased track inclination angle i m and increase the forward inclination angle i n Determining the inclination of the reference orbit includes:

[0118] Step S1: Set the first track inclination angle i1 and the second track inclination angle i2; the initial value of the first track inclination angle i1 is the track inclination angle i2 before increasing the inclination angle i2. n The initial value of the second orbital inclination angle i2 is the orbital inclination angle i after it is increased. m ;

[0119] Step S2: Determine the second current orbital inclination i′ as the median of the first and second orbital inclinations, i′ = (i1 + i2) / 2; Under the second current orbital inclination, determine the first time required for the satellite to travel from the first latitude argument to the second latitude argument and the second time required for the satellite to travel from the geographical longitude of the first ascending node to the geographical longitude of the second ascending node; Determine the time difference between the first and second times under the second current orbital inclination; Determine whether the absolute value of the time difference between the first and second times under the second current orbital inclination is less than a preset time threshold. If it is less, then determine the second current orbital inclination as the reference orbital inclination.

[0120] Step S3: If the absolute value of the time difference between the first time and the second time under the second current track inclination is not less than the time threshold, then determine the track inclination angle with the opposite sign to the time difference under the second current track inclination angle i1 and the second track inclination angle i2; reassign the two track inclination angles with the opposite sign to the first track inclination angle i1 and the second track inclination angle i2.

[0121] Repeat steps S2 and S3 until the absolute value of the time difference corresponding to the second current orbital inclination is less than the preset time threshold. Then, determine the second current orbital inclination whose absolute value of the time difference is less than the time threshold as the reference orbital inclination in the combined mode.

[0122] Continuing with the example above, suppose that in the 8th calculation, Δt is 10s, and the corresponding orbital inclination angle is 30°. Since |Δt| is not less than 0.1s, a 9th calculation will be performed.

[0123] In the 9th calculation, the track inclination angle is increased by 1°, that is, the increased track inclination angle is 31°. Assuming that the Δt obtained in the 9th calculation is 8s, since the Δt corresponding to the 9th calculation and the Δt corresponding to the 8th calculation are both positive values ​​(same sign), the track inclination angle corresponding to the 9th calculation will be increased by another 1°, that is, the increased track inclination angle is 32°, and then the 10th calculation is performed.

[0124] Assuming the 10th calculation yields a Δt of -0.8s, and since the Δt values ​​from the 9th and 10th calculations have opposite signs, the orbital inclination angle satisfying |Δt| < 0.1s must be between 31° and 32°. Therefore, i is determined... n It is 31°, i m It is 32°.

[0125] Set a first track inclination angle i1 and a second track inclination angle i2. The initial value of the first track inclination angle i1 is 31° and the initial value of the second track inclination angle i2 is 32°.

[0126] The second current orbital inclination angle i′ is set, and its initial value is (31+32) / 2 = 31.5°. The 11th calculation is performed with 31.5° as the reference. Assuming that the 11th calculation yields 0.07s of Δt, since |Δt| is less than 0.1s, the condition is met. Therefore, 31.5° will be determined as the reference orbital inclination angle, and the first duration t corresponding to 31.5° will be obtained. u According to formula T w =T0-t u Determine the corresponding launch window.

[0127] Assuming the Δt obtained in the 11th calculation is 0.7s, since |Δt| is not less than 0.1s, a 12th calculation will be performed.

[0128] When performing the 12th calculation, it is necessary to find the orbital inclination angle that corresponds to the time difference under the second current orbital inclination angle with the opposite sign among the first orbital inclination angle i1 and the second orbital inclination angle i2 used in the 11th calculation, and reassign these two orbital inclination angles to the first orbital inclination angle i1 and the second orbital inclination angle i2.

[0129] For example, in the 11th calculation, the first inclination angle i1 is 31°, the second orbital inclination angle i2 is 32°, and the second current orbital inclination angle i′ is 31.5°. Since the duration difference corresponding to 31° is 8s, the duration difference corresponding to 32° is -0.08s, and the duration difference corresponding to 31.5° is 0.07s, the orbital inclination angle corresponding to the duration difference with the opposite sign to 0.07 is 32°. Therefore, 31.5° will be reassigned to the first orbital inclination angle i1, and 32° will be reassigned to the second orbital inclination angle i2. Then, step S2 or S3 will be executed again to finally determine the second current orbital inclination angle whose absolute value of the duration difference is less than the duration threshold. This second orbital inclination angle is then determined as the reference orbital inclination angle in the combined mode.

[0130] It should be noted that if, during the iteration process of any combination mode, the absolute value of the duration difference is not less than the preset duration threshold or the duration difference shows an opposite sign even when the orbital inclination angle is greater than π-B_m, then it is considered that there is no solution for that combination mode.

[0131] This determines the satellite's reference orbital inclination and corresponding launch window for this combined mode.

[0132] Using the same method described above, the corresponding reference orbit inclination and launch window for all combination modes can be determined. Finally, the minimum orbit inclination among all reference orbit inclinations is determined as the target orbit inclination of the satellite, and the reference launch window corresponding to the target orbit inclination is taken as the target launch window.

[0133] This determines the target orbital inclination and corresponding launch window for a satellite when the launch point is fixed and the orbital inclination is unknown. The launch window is the time range within which the satellite can be launched.

[0134] S112, if it is determined that the launch point of the satellite is an unknown value and the orbital inclination is a fixed value, the longitude and latitude of the nadir point corresponding to each launch window are determined according to the coordinates of the observation point, the target observation time, and the preset observation time step.

[0135] When the satellite's launch point is unknown and its orbital inclination is fixed, and the target observation time is pre-set while the current time is known, multiple observable time points (launch windows) can be inferred from the target observation time based on the observation time step. Then, the longitude and latitude of the nadir point for each launch window can be determined. The specific implementation is as follows:

[0136] For any combination pattern among all combination patterns, perform the following processing:

[0137] Obtain the predetermined orbital inclination of the satellite, and determine the third latitude argument and the geographical longitude of the third ascending node when the satellite passes through the observation point based on the predetermined orbital inclination; wherein, the calculation formula for the third latitude argument can refer to the above formula (5) or formula (7); the geographical longitude of the third ascending node can refer to the above formula (6) and formula (8) to determine.

[0138] According to formula (17), determine the geographical longitude Ω of the current ascending node corresponding to the satellite's observation time n observation time steps Δ′t from the target observation time. t ;

[0139] Ω t =Ω2+(W Ω ′-ω e )×nΔ′t (17)

[0140] According to formula (18), determine the current latitude argument u corresponding to the satellite regressing n observation time steps Δ′t from the target observation time. t :

[0141] u t =u2+(W′) ω +W′ M )×nΔ′t (18)

[0142] Determine the geographic latitude of the nadir point when the satellite regresses n observation time steps from the target observation time based on the current latitude argument;

[0143] The geographic longitude of the nadir point is determined based on the current ascending node's geographic longitude, counting backwards n observation time steps from the target observation time; where...

[0144] Ω2 is the third latitude argument, u2 is the geographical longitude of the third ascending node, and W′ Ω ω is the precession rate of the right ascension of the ascending node of the satellite at a preset orbital inclination. e Let W′ be the Earth's rotational angular velocity. ω The perigee precession rate of the satellite at the preset orbital inclination, W′ M This is the precession rate of the satellite at its mean apogee angle under a preset orbital inclination.

[0145] In one implementation, determining the geographic latitude of the nadir point when the satellite regresses n observation time steps from the target observation time based on the current latitude argument includes:

[0146] The geographic latitude B of the nadir point when the satellite regresses n observation time steps from the observation time of the target is determined according to formula (19). t :

[0147] B t =arcsin(sin(u t )×sin(i″)) (19)

[0148] Among them, u t Let i″ be the current latitude argument when the satellite moves back n observation time steps Δ′t from the target observation time, and let i″ be the preset orbital inclination angle.

[0149] Based on the current ascending node longitude, determine the nadir longitude of the satellite when it regresses n observation time steps from the target observation time, including:

[0150] According to the formula Determine the geographic longitude L of the nadir point when the satellite regresses n observation time steps from the target observation time. t ;in,

[0151] The Ω t The geographical longitude of the current ascending node of the satellite is defined as n observation time steps Δ′t, which is the time step backward from the observation time of the target. The preset orbital inclination is defined as u. t The current latitude argument is the angle at which the satellite moves back n observation time steps Δ′t from the observation time of the target.

[0152] The minimum value of n is 1, and the maximum value is determined based on the current time, the preset observation time, and the observation step size.

[0153] For example, if the current time is 9:00, the preset observation time is 21:00 on the same day, and the preset observation step size is 1 hour, then the maximum value of n is 12. Starting from 21:00, the observation window will be rolled back every hour, resulting in 12 launch windows.

[0154] In each combination mode (in this case, since the launch point is unknown, only two combination modes are included: the mode where the satellite ascends to pass the observation point with an increasing orbital inclination and the mode where the satellite descends to pass the observation point with an increasing orbital inclination), 12 launch windows and 12 launch point coordinates (nadir latitude and nadir longitude) can be determined. These two combination modes together determine 24 launch windows and their corresponding 24 launch point coordinates. In actual launches, the most suitable launch window and corresponding launch point coordinates can be determined as needed.

[0155] In this invention, in emergency launch scenarios, a launch window can be determined according to launch requirements (fixed orbital inclination or fixed launch point). Then, based on the launch window and launch point, the satellite can be launched quickly and accurately to a preset position to achieve fixed-point observation of emergency events.

[0156] Based on the same inventive concept as in the foregoing embodiments, this embodiment also provides a device for launching a satellite in an emergency scenario, such as... Figure 2 As shown, the device includes:

[0157] The first determining unit 21 is used to determine, if the launch point of the satellite is determined to be a fixed value and the orbital inclination is an unknown value, the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point under the first current orbital inclination, based on the coordinates of the observation point and the coordinates of the launch point; and to determine, based on the coordinates of the observation point and the coordinates of the launch point, the second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point under the first current orbital inclination.

[0158] The second determining unit 22 is used to determine the target orbital inclination of the satellite and the corresponding launch window based on the first latitude argument, the first ascending node longitude, the second latitude argument, and the second ascending node longitude.

[0159] The third determining unit 23 is used to determine the longitude and latitude of the nadir point corresponding to each launch window if the launch point of the satellite is an unknown value and the orbital inclination is a fixed value, based on the coordinates of the observation point, the target observation time, and the preset observation time step.

[0160] Since the apparatus described in the embodiments of this invention is used to implement the method of launching satellites in emergency scenarios according to the embodiments of this invention, those skilled in the art can understand the specific structure and variations of the apparatus based on the method described in the embodiments of this invention, and therefore will not be described in detail here. All apparatuses used in the methods of the embodiments of this invention fall within the scope of protection of this invention.

[0161] Through one or more embodiments of the present invention, the present invention has the following beneficial effects or advantages:

[0162] This invention provides a method, apparatus, and medium for launching a satellite based on an emergency scenario. The method includes: if the launch point of the satellite is determined to be a fixed value and the orbital inclination is an unknown value, determining, based on the coordinates of the observation point and the coordinates of the launch point, the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point at a first current orbital inclination; and determining, based on the coordinates of the observation point and the coordinates of the launch point, the second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point at the first current orbital inclination; and determining, based on the first latitude argument, the geographical longitude of the first ascending node, and the coordinates of the launch point, the second latitude argument and the second ascending node geographical longitude corresponding to the satellite passing through the observation point at the first current orbital inclination; and determining, based on the first latitude argument, the first ascending node geographical longitude, and the second ascending node geographical longitude, the satellite launch point is launched at an emergency scenario. The target orbital inclination of the satellite and the corresponding launch window are determined by the second latitude argument and the geographic longitude of the second ascending node. If the launch point of the satellite is unknown and the orbital inclination is fixed, the longitude and latitude of the nadir point under each launch window are determined according to the coordinates of the observation point, the target observation time, and the preset observation time step. In this way, in emergency launch scenarios, the launch window can be determined according to the launch requirements (fixed orbital inclination or fixed launch point), and the satellite can be launched quickly and accurately to the preset position based on the launch window and launch point to achieve fixed-point observation of emergency events.

[0163] Although preferred embodiments of the invention have been described, those skilled in the art, upon learning the basic inventive concept, can make other changes and modifications to these embodiments. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments as well as all changes and modifications falling within the scope of the invention.

[0164] The above description is merely a preferred embodiment of the present invention and is not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for launching satellites based on emergency scenarios, characterized in that, The method includes: If the launch point of the satellite is determined to be a fixed value and the orbital inclination is an unknown value, the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point under the first current orbital inclination are determined according to the coordinates of the observation point and the coordinates of the launch point. The second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point under the first current orbital inclination are determined according to the coordinates of the observation point and the coordinates of the launch point. The target orbital inclination of the satellite and the corresponding launch window are determined based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node. If the launch point of the satellite is determined to be an unknown value and the orbital inclination is a fixed value, the longitude and latitude of the nadir point under each launch window are determined according to the coordinates of the observation point, the target observation time, and the preset observation time step.

2. The method as described in claim 1, characterized in that, The first latitude argument includes: a first sub-latitude argument and a second sub-latitude argument; the first ascending node longitude includes: a first sub-ascending node longitude and a second sub-ascending node longitude; determining the first latitude argument and the first ascending node longitude corresponding to the launch point at the first current orbital inclination, based on the observation point coordinates and the launch point coordinates, includes: If it is determined that the satellite's orbital descent passes through the launch point, according to the formula... Determine the first sub-latitude argument u0 corresponding to the satellite's descent orbit passing the launch point at the first current orbital inclination angle; The geographical longitude Ω0 of the first sub-ascending node corresponding to the satellite when the first current orbital inclination descends through the launch point is determined according to the formula Ω0=L0-arctan(cos(i)·tan(u0))-π. If it is determined that the satellite's orbital ascent passes through the launch point, according to the formula... Determine the second sub-latitude argument u′0 corresponding to the satellite's ascent and passage through the launch point at the first current orbital inclination; The geographical longitude Ω′0 of the second sub-ascending node corresponding to the satellite's ascent at the launch point under the first current orbital inclination is determined according to the formula Ω′0=L0-arctan(cos(i)·tan(u0)); where Ω′0=L0-arctan(cos(i)·tan(u0)) B0 is the latitude of the launch point, i is the first current orbital inclination, i∈[B_m,π-B_m], B_m=max(|B0|,|B1|), B1 is the latitude of the observation point, and L0 is the longitude of the launch point.

3. The method as described in claim 1, characterized in that, The second latitude argument includes: a third sub-latitude argument and a fourth sub-latitude argument; the second ascending node longitude includes: a third sub-ascending node longitude and a fourth sub-ascending node longitude; determining the second latitude argument and the second ascending node longitude corresponding to the satellite passing the observation point at the first current orbital inclination includes: If it is determined that the satellite's orbital descent passes through the observation point, according to the formula... Determine the third sub-latitude argument u1 corresponding to when the satellite's descending orbit at the first current orbital inclination passes through the observation point; The geographical longitude Ω1 of the third sub-ascending node corresponding to the observation point when the satellite's descending orbit at the first current orbital inclination passes through the observation point is determined according to the formula Ω1=L1-arctan(cos(i)·tan(u1))-π. If it is determined that the satellite's orbital ascent passes through the observation point, according to the formula... Determine the fourth sub-latitude argument u′1 corresponding to the observation point when the satellite ascends to its orbit at the first current orbital inclination. The geographical longitude Ω′1 of the fourth sub-ascending node corresponding to the observation point when the satellite ascends to its first current orbital inclination is determined according to the formula Ω′1=L1-arctan(cos(i)·tan(u1)); where Ω′1=L1-arctan(cos(i)·tan(u1)) The i is the first current orbital inclination angle, the i ∈ [B_m, π-B_m], the B_m = max(|B0|, |B1|), the B1 is the latitude of the observation point, the B0 is the latitude of the launch point, and the L1 is the longitude of the observation point.

4. The method as described in claim 1, characterized in that, The process of determining the target orbital inclination of the satellite and the corresponding launch window based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node includes: For any combination pattern among all combination patterns, perform the following processing: Determine the precession rate W of the right ascension of the satellite's ascending node at the first current orbital inclination. Ω Perimeter angle precession rate W ω and the precession rate W at the approximate point angle M According to the formula Determine the first time t required for the satellite to travel from the first latitude angle to the second latitude angle. u According to the formula The second time t required for the satellite to precess from the first ascending node's geographic longitude to the second ascending node's geographic longitude is determined. Ω If the absolute value of the duration difference between the first duration and the second duration is determined to be less than a preset duration threshold, then the first current orbital inclination angle is determined as the reference orbital inclination angle in the combined mode; and T0-t u The reference launch window is determined as the reference orbit inclination angle in the combined mode; the combined mode includes: the mode in which the satellite descends to a lower orbit and passes through the launch point and the observation point, the mode in which the satellite ascends to a higher orbit and passes through the launch point and the observation point, the mode in which the satellite ascends to a higher orbit and passes through the launch point and the satellite descends to a lower orbit and the observation point, and the mode in which the satellite descends to a lower orbit and passes through the launch point and the satellite ascends to a higher orbit and passes through the observation point; The minimum value among all reference orbit inclinations obtained under all combination modes is determined as the target orbit inclination of the satellite, and the reference launch window corresponding to the target orbit inclination is the target launch window; Wherein, Δu is the difference in latitude argument between the first latitude argument and the second latitude argument, ΔΩ is the difference in geographical longitude between the first ascending node and the second ascending node, and T0 is the time when the satellite observes the observation point.

5. The method as described in claim 4, characterized in that, If the absolute value of the duration difference between the first duration and the second duration is not less than a preset duration threshold, then the orbital inclination angle is increased according to the orbital inclination angle step size. Each time it is increased, the following processing needs to be performed: Under the increased orbital inclination, determine the first time required for the satellite to travel from the first latitude argument to the second latitude argument, and the second time required for the satellite to travel from the geographical longitude of the first ascending node to the geographical longitude of the second ascending node; determine the time difference between the first time and the second time under the increased orbital inclination; further determine whether the absolute value of the time difference under the increased orbital inclination is less than the time threshold. If it is less, then the increased orbital inclination is determined as the reference orbital inclination in the combined mode. If it is not less than, then determine whether the time difference under the increased track inclination angle has the same sign as the time difference under the original track inclination angle. If the signs are the same, continue to increase the track inclination angle and repeat the above steps; repeat this process until the increased track inclination angle i is determined. m The difference in duration and the increase in the orbital inclination angle i n If the time difference values ​​are of opposite signs, then the inclination angle of the reference orbit is determined to be within [i]. n i m Within the interval, according to the increased track inclination angle i m and the increased front track inclination angle i n Determine the inclination angle of the reference orbit.

6. The method as described in claim 5, characterized in that, According to the increased track inclination angle i m and the increased front track inclination angle i n Determining the inclination angle of the reference orbit includes: Step S1: Set the first track inclination angle i1 and the second track inclination angle i2; the initial value of the first track inclination angle i1 is the track inclination angle i2 before the increase. n The initial value of the second track inclination angle i2 is the increased track inclination angle i. m ; Step S2: Determine the second current orbital inclination i′ as the median of the first orbital inclination and the second orbital inclination, i′ = (i1 + i2) / 2; Under the second current orbital inclination, determine the first time required for the satellite to travel from the first latitude argument to the second latitude argument and the second time required for the satellite to travel from the geographical longitude of the first ascending node to the geographical longitude of the second ascending node; Determine the time difference between the first time and the second time under the second current orbital inclination; Determine whether the absolute value of the time difference between the first time and the second time under the second current orbital inclination is less than a preset time threshold. If it is less, then determine the second current orbital inclination as the reference orbital inclination. Step S3: If the absolute value of the time difference between the first time and the second time under the second current track inclination is not less than the time threshold, then determine the track inclination angle that corresponds to the time difference under the second current track inclination angle with the opposite sign from the first track inclination angle i1 and the second track inclination angle i2; reassign the two track inclination angles to the first track inclination angle i1 and the second track inclination angle i2. Repeat steps S2 and S3 until the absolute value of the time difference corresponding to the second current orbital inclination is less than the preset time threshold. Then, determine the second current orbital inclination whose absolute value of the time difference is less than the time threshold as the reference orbital inclination in the combined mode.

7. The method as described in claim 1, characterized in that, The process of determining the longitude and latitude of the nadir point corresponding to each launch window based on the observation point coordinates, the target observation time, and the preset observation time step includes: For any combination pattern among all combination patterns, perform the following processing: Obtain the predetermined orbital inclination of the satellite, and determine the third latitude argument and the geographical longitude of the third ascending node when the satellite passes through the observation point based on the predetermined orbital inclination; According to the formula Ω t =Ω2+(W′) Ω -ω e The geographic longitude Ω of the current ascending node is determined by )×nΔ′t when the satellite moves back n observation time steps Δ′t from the observation time of the target. t ; According to formula u t =u2+(W′) ω +W′ M The current latitude argument u is determined by )×nΔ′t when the satellite moves back n observation time steps Δ′t from the observation time of the target. t ; Determine the geographic latitude of the nadir point when the satellite regresses n observation time steps from the target observation time based on the current latitude argument; Based on the current ascending node's geographic longitude, determine the nadir longitude of the satellite when it regresses n observation time steps from the target observation time; where... Ω2 is the third latitude argument, u2 is the geographical longitude of the third ascending node, and W′ is... Ω ω is the precession rate of the right ascension of the ascending node of the satellite at the preset orbital inclination. e The angular velocity of Earth's rotation, W′ ω W′ is the perigee argument precession rate of the satellite at the preset orbital inclination. M The precession rate of the satellite at the mean apogee angle under the preset orbital inclination.

8. The method as described in claim 7, characterized in that, The step of determining the geographic latitude of the nadir point when the satellite regresses n observation time steps from the target observation time based on the current latitude argument includes: According to formula B t =arcsin(sin(u t The geographic latitude B of the nadir point is determined by the function )×sin(i″)) when the satellite regresses n observation time steps from the observation time of the target. t ;in, The u t The current latitude argument is the angle at which the satellite moves backward n time steps from the observation time of the target, and i″ is the preset orbital inclination angle.

9. The method as described in claim 7, characterized in that, The step of determining the nadir longitude of the satellite when it has regressed n observation time steps from the target observation time based on the current ascending node longitude includes: According to the formula Determine the geographic longitude L of the nadir point when the satellite regresses n observation time steps from the target observation time. t ;in, The Ω t The geographical longitude of the current ascending node of the satellite is defined as n observation time steps Δ′t, which is the time step backward from the observation time of the target. The preset orbital inclination is defined as u. t The current latitude argument is the angle at which the satellite moves back n observation time steps Δ′t from the observation time of the target.

10. A device for launching satellites based on emergency scenarios, characterized in that, The device includes: The first determining unit is configured to, if the launch point of the satellite is determined to be a fixed value and the orbital inclination is an unknown value, determine the first latitude argument and the geographical longitude of the first ascending node corresponding to the satellite passing through the launch point under the first current orbital inclination, based on the coordinates of the observation point and the coordinates of the launch point; and determine the second latitude argument and the geographical longitude of the second ascending node corresponding to the satellite passing through the observation point under the first current orbital inclination, based on the coordinates of the observation point and the coordinates of the launch point. The second determining unit is used to determine the target orbital inclination of the satellite and the corresponding launch window based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node; The third determining unit is used to determine the longitude and latitude of the nadir point corresponding to each launch window if the launch point of the satellite is an unknown value and the orbital inclination is a fixed value, based on the coordinates of the observation point, the target observation time, and the preset observation time step.

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