A method for planning detailed investigation missions of geosynchronous orbit target imaging satellites

Abstract

The present invention discloses a method for planning a detailed inspection mission of a geosynchronous orbit target imaging satellite, the method comprising: dividing the mission orbit into an approaching segment, a phase adjustment segment, an outbound detailed inspection segment, a U-turn segment, a return detailed inspection segment, and a distance segment; determining the maneuvering time of the first orbital maneuver, the second orbital maneuver, the third orbital maneuver, the fourth orbital maneuver, and the fifth orbital maneuver; determining the speed increment of the first orbital maneuver, the second orbital maneuver, the third orbital maneuver, the fourth orbital maneuver, and the fifth orbital maneuver. The present invention does not have a long-term approaching state with the target, which is conducive to concealing the observation intention, reducing the intensity of the on-orbit game, avoiding legal disputes, and at the same time avoiding the target's awareness and evasion of observation as much as possible, which is conducive to achieving the observation purpose.

Description

Technical Field

[0001] The present invention belongs to the technical field of space situation awareness space target monitoring imaging satellite mission planning, and in particular relates to a method for planning a key detailed inspection mission of a geosynchronous orbit target imaging satellite. Background Art

[0002] The mission modes of space target surveillance and imaging satellites generally include drift patrol survey and key target detailed survey. Drift patrol survey is one of the commonly used mission modes. The satellite orbit is located below the target orbit. The orbit height difference forms the drift rate difference between the satellite and the target. The satellite completes surveillance imaging of the target when drifting below the target. The satellite usually uses more than 100 days to complete drift patrol of multiple targets in the mission area. It is a commonly used survey method.

[0003] As the space target imaging satellite spends more time in orbit, after completing the drift inspection of the targets in the mission area, it will switch to the detailed inspection mission mode, that is, it will stay in a certain orbit for a long time. When a target orbit maneuvers, a new target enters orbit, or various space events (approach, docking, etc.) occur in the mission area, it will quickly drift to the vicinity of the target to complete a detailed inspection of the target. After the detailed inspection is completed, it will drift back to the original orbit to continue control. The detailed inspection mission is a mission mode that space target imaging satellites will use for a long time in the middle and late stages of their orbit.

[0004] At present, detailed inspection missions generally adopt the method of drifting-flying-flying around-drifting back, that is, after drifting to the vicinity of the target, it will first be fixed near the target, then choose the appropriate time to start the detailed inspection by flying around in the direction of the light, and finally exit the flying around and drift back to the original orbit. This method has the following limitations: (1) The mission concealment is poor, and there is a long period of close proximity to the target, which is easy to cause the target to counter-reconnaissance maneuvers to avoid observation, affecting intelligence acquisition, and the game intensity is strong, which has the risk of causing legal disputes; (2) The detailed inspection time is concentrated in the flying around stage, which is easy for the target to avoid reconnaissance, and the degree of interference with the target mission is limited; (3) The speed increment consumption is large, which affects the execution efficiency of the entire satellite mission. The speed increment consumption for completing a detailed inspection mission for a single target is generally about 10m / s. Summary of the invention

[0005] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and provide a method for planning a key detailed inspection mission for a geosynchronous orbit target imaging satellite, in which there is no long-term proximity to the target, which is conducive to concealing the observation intention, reducing the intensity of on-orbit gaming, and avoiding legal disputes. At the same time, it is possible to avoid causing the target to notice and evade observation, which is conducive to achieving the observation purpose.

[0006] The object of the present invention is achieved through the following technical solutions: a method for planning a detailed investigation mission of a geosynchronous orbit target imaging satellite, comprising: dividing the mission orbit into an approaching segment, a phasing segment, an outbound detailed investigation segment, a U-turn segment, a return detailed investigation segment and a distance segment; wherein a first orbital maneuver is performed between the approaching segment and the phasing segment, a second orbital maneuver is performed between the phasing segment and the outbound detailed investigation segment, a third orbital maneuver is performed between the outbound detailed investigation segment and the U-turn segment, a fourth orbital maneuver is performed between the U-turn segment and the return detailed investigation segment, and a fifth orbital maneuver is performed between the return detailed investigation segment and the distance segment; determining the maneuvering times of the first orbital maneuver, the second orbital maneuver, the third orbital maneuver, the fourth orbital maneuver and the fifth orbital maneuver; determining the speed increments of the first orbital maneuver, the second orbital maneuver, the third orbital maneuver, the fourth orbital maneuver and the fifth orbital maneuver.

[0007] In the above-mentioned geosynchronous orbit target imaging satellite focused detailed investigation mission planning method, determining the maneuvering time T2 of the second orbit maneuver includes determining the time and determining the date; wherein the determined time is 12:00 noon local time at the geographical longitude L0 of the preset target subsatellite point; determining the date includes: if the time T0 of our satellite at the initial position is earlier than 12:00 noon local time at our subsatellite point, then the maneuvering time of the second orbit maneuver is If the time of T0 is later than 12:00 noon when the star is below us, then Where R0 is the semi-major axis of the geosynchronous orbit, a0 is the semi-major axis, and ceiling means the result is rounded up.

[0008] In the above-mentioned geosynchronous orbit target imaging satellite focused detailed investigation mission planning method, the first orbital maneuvering time is T1=T2-24h; wherein T1 is the first orbital maneuvering time, and T2 is the maneuvering time of the second orbital maneuvering.

[0009] In the above-mentioned geosynchronous orbit target imaging satellite key detailed investigation mission planning method, the third orbital maneuvering time is T3=T2+n·24h; wherein T3 is the third orbital maneuvering time, and n is a positive integer greater than 1.

[0010] In the above-mentioned geosynchronous orbit target imaging satellite key detailed investigation mission planning method, the fourth orbital maneuvering time is T4=T3+12h; wherein T3 is the third orbital maneuvering time, and T4 is the fourth orbital maneuvering time.

[0011] In the above-mentioned geosynchronous orbit target imaging satellite key detailed investigation mission planning method, the fifth orbital maneuvering time is T5=T4+n·24h; wherein T5 is the fifth orbital maneuvering time, T4 is the fourth orbital maneuvering time, and n is a positive integer greater than 1.

[0012] In the above-mentioned geosynchronous orbit target imaging satellite focused detailed inspection mission planning method, determining the velocity increment of the first orbital maneuver and the velocity increment of the second orbital maneuver includes: determining the initial value Δv of the velocity increment of the second orbital maneuver 2_0 , so that the outbound detailed inspection segment orbit crosses the geosynchronous orbit twice a day; the distance between the easternmost crossing point of the T2~T2+24h segment and the westernmost crossing point of the T2+24h~T2+48h segment is equal to 2D0, and the rough optimization intermediate value Δv of the velocity increment of the second orbital maneuver is obtained 2_1 ; Where D0 is the expected observation distance for detailed inspection of the target; when T2+24h and In parallel, we obtain the velocity increment of the first orbital maneuver and the velocity increment of the second orbital maneuver; where, is the position vector of the Earth's center pointing to our star at T2+24h, It is the position vector pointing from the center of the earth to the target at T2+24h.

[0013] In the above-mentioned geosynchronous orbit target imaging satellite focused detailed investigation mission planning method, the velocity increment Δv3 of the third orbital maneuver is obtained by the following formula:

[0014]

[0015] Where a0 is the semi-major axis, Δv1 is the velocity increment of the first orbital maneuver, Δv2 is the velocity increment of the second orbital maneuver, μ is the gravitational constant at the center of the Earth, and R0 is the semi-major axis of the geosynchronous orbit.

[0016] In the above-mentioned geosynchronous orbit target imaging satellite focused detailed inspection mission planning method, determining the velocity increment of the fourth orbital maneuver includes: determining the initial value Δv of the velocity increment of the fourth orbital maneuver 4_0 , so that the return detailed inspection section orbit crosses the geosynchronous orbit twice a day; when T5-24h time and In parallel, we get the velocity increment of the fourth orbital maneuver; where, is the position vector of the Earth's center pointing to our star at T5-24h, It is the position vector of the center of the earth pointing to the target at T5-24h.

[0017] In the above-mentioned geosynchronous orbit target imaging satellite focused detailed investigation mission planning method, the velocity increment Δv5 of the fifth orbital maneuver is obtained by the following formula:

[0018]

[0019] Where a0 is the semi-major axis, Δv4 is the velocity increment of the second orbital maneuver, μ is the gravitational constant at the center of the Earth, and R0 is the semi-major axis of the geosynchronous orbit.

[0020] Compared with the prior art, the present invention has the following beneficial effects:

[0021] (1) The observation concealment of the present invention is strong. The satellite is always in a drifting state. Compared with the currently commonly used accompanying and circling strategies, there is no long-term close proximity to the target, which is conducive to concealing the observation intention, reducing the intensity of on-orbit game, and avoiding legal disputes. At the same time, it is possible to avoid causing the target to notice and evade observation, which is conducive to achieving the observation purpose;

[0022] (2) The present invention can conduct four close-in detailed inspections from four different directions (west side facing the ground, west side facing away from the ground, east side facing the ground, east side facing away from the ground) within a few days, all in the condition of direct sunlight. In addition, it can focus on observing the east and west sides of the target (conventional drift patrol can only observe a single target once, and focus on observing the target facing the ground and away from the ground, while the key payloads of most space situational awareness and space attack and defense satellites are located on the east and west sides). The close-in reconnaissance process is dispersed within a few days, making it more difficult and costly for the target to avoid reconnaissance, and the observation results have higher intelligence value;

[0023] (3) When the present invention completes the observation of the same target under the same conditions, compared with the currently commonly used accompanying and circling strategies, the speed increment consumption is reduced by more than 30%, which can greatly improve the reconnaissance efficiency of the entire satellite. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Various other advantages and benefits will become apparent to those of ordinary skill in the art by reading the detailed description of the preferred embodiments below. The accompanying drawings are only for the purpose of illustrating the preferred embodiments and are not to be considered as limiting the present invention. Moreover, the same reference symbols are used throughout the accompanying drawings to represent the same components. In the accompanying drawings:

[0025] Figure 1 It is a flow chart of a method for planning a geosynchronous orbit target imaging satellite focused detailed inspection mission provided by an embodiment of the present invention;

[0026] Figure 2 It is a schematic diagram of the relative motion relationship between the sun, the earth, a satellite and a target provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0027] Exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although exemplary embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided in order to enable a more thorough understanding of the present disclosure and to be able to fully convey the scope of the present disclosure to those skilled in the art. It should be noted that, in the absence of conflict, the embodiments of the present invention and the features described in the embodiments can be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0028] Figure 1 1 is a flow chart of a method for planning a geosynchronous orbit target imaging satellite focused detailed inspection mission provided by an embodiment of the present invention. Figure 1 As shown, the geosynchronous orbit target imaging satellite focused detailed inspection mission planning method includes:

[0029] The mission track is divided into an approaching section, a phasing section, an outbound detailed inspection section, a U-turn section, a return detailed inspection section, and a distance section; wherein the first orbital maneuver is performed between the approaching section and the phasing section, the second orbital maneuver is performed between the phasing section and the outbound detailed inspection section, the third orbital maneuver is performed between the outbound detailed inspection section and the U-turn section, the fourth orbital maneuver is performed between the U-turn section and the return detailed inspection section, and the fifth orbital maneuver is performed between the return detailed inspection section and the distance section;

[0030] determining maneuver times for a first orbital maneuver, a second orbital maneuver, a third orbital maneuver, a fourth orbital maneuver, and a fifth orbital maneuver;

[0031] Determine the velocity increments for the first orbital maneuver, the second orbital maneuver, the third orbital maneuver, the fourth orbital maneuver, and the fifth orbital maneuver.

[0032] Determining the maneuvering time T2 of the second orbital maneuver includes determining the time and determining the date; wherein,

[0033] The time is determined to be 12:00 noon local time at the geographical longitude L0 of the preset target sub-satellite point;

[0034] The date to be determined includes: If the time T0 of our satellite at the initial position is earlier than 12:00 noon local time of our satellite's lower point, then the maneuvering time of the second orbital maneuver is If the time of T0 is later than 12:00 noon when the star is below us, then Where R0 is the semi-major axis of the geosynchronous orbit, and ceiling means the result is rounded upwards.

[0035] The first orbital maneuvering time is T1=T2-24h; wherein T1 is the first orbital maneuvering time, and T2 is the second orbital maneuvering time.

[0036] The third orbital maneuvering time is T3=T2+n·24h; wherein T3 is the third orbital maneuvering time, and n is a positive integer greater than 1.

[0037] The fourth orbital maneuvering time is T4=T3+12h; wherein T3 is the third orbital maneuvering time, and T4 is the fourth orbital maneuvering time.

[0038] The fifth orbital maneuvering time is T5=T4+n·24h; wherein T5 is the fifth orbital maneuvering time, T4 is the fourth orbital maneuvering time, and n is a positive integer greater than 1.

[0039] Determining the velocity increment for the first orbital maneuver and the velocity increment for the second orbital maneuver includes:

[0040] Determine the initial value of the velocity increment Δv for the second orbital maneuver 2_0 , so that the outbound detailed inspection segment orbit crosses the geosynchronous orbit belt twice a day;

[0041] Make the distance between the easternmost crossing point of the T2~T2+24h segment and the westernmost crossing point of the T2+24h~T2+48h segment equal to 2D0, and obtain the rough optimization intermediate value Δv2 of the velocity increment of the second orbital maneuver; where D0 is the expected observation distance for detailed inspection of the target;

[0042] When T2+24h and In parallel, we obtain the velocity increment of the first orbital maneuver and the velocity increment of the second orbital maneuver; where, is the position vector of the Earth's center pointing to our star at T2+24h, It is the position vector pointing from the center of the earth to the target at T2+24h.

[0043] The velocity increment Δv3 of the third orbital maneuver is obtained by the following formula:

[0044]

[0045] Where a0 is the semi-major axis, Δv1 is the velocity increment of the first orbital maneuver, Δv2 is the velocity increment of the second orbital maneuver, and μ is the gravitational constant at the center of the Earth (μ = 398600.44 km 3 / s 2 ), R0 is the semi-major axis of the geosynchronous orbit.

[0046] The velocity increments for the fourth orbital maneuver are determined by:

[0047] Determine the initial value of the velocity increment Δv for the fourth orbital maneuver 4_0, so that the return detailed inspection segment orbit crosses the geosynchronous orbit belt twice a day;

[0048] When T5-24h and In parallel, we get the velocity increment of the fourth orbital maneuver; where, is the position vector of the Earth's center pointing to our star at T5-24h, is the position vector pointing from the center of the earth to the target at that moment.

[0049] The velocity increment Δv5 of the fifth orbital maneuver is obtained by the following formula:

[0050]

[0051] Where a0 is the semi-major axis, Δv4 is the velocity increment of the second orbital maneuver, and μ is the gravitational constant at the center of the Earth (μ = 398600.44 km 3 / s 2 ), R0 is the semi-major axis of the geosynchronous orbit.

[0052] This embodiment uses a "water drop" shaped track with a large eccentricity. Without performing high-intensity games such as accompanying flight and circling flight, it uses six tracks to complete four close-up observations of the target in a highly concealed, forward-lighting, and multi-directional manner. The steps are as shown in the attached figure. Figure 1 shown.

[0053] The specific steps are:

[0054] (1) Detailed inspection mission track segment design. The mission planning method proposed in the present invention divides the entire detailed inspection mission track into 6 parts, namely, the approach segment, the phase adjustment segment, the outbound detailed inspection segment, the U-turn segment, the return detailed inspection segment, and the outbound segment. One track maneuver is performed between every two track segments, for a total of 5 track maneuvers. Figure 1 As shown. There are two opportunities for close inspection in the outbound detailed inspection section and the return detailed inspection section, a total of four times, all in the condition of direct sunlight, and two detailed inspections can be carried out on the east and west sides of the target respectively. At the same time, it can ensure that when the target is in direct sunlight to our satellite, our satellite will keep a long distance from the target, reducing the risk of our satellite being observed by the target.

[0055] During the approach phase, our satellite is in a fast-drifting circular orbit. In order to ensure a certain drift rate and to keep the distance from the target as large as possible when backlit, the general orbit height difference with the GEO orbit height is ≥80km;

[0056] Our satellite completed the first maneuver at position 1, with a velocity increment of Δv1, and entered the phase adjustment phase. The purpose of phase adjustment is to ensure that the following four close-in observations are all in the straight light condition;

[0057] Our satellite completed the second maneuver at position 2, with a speed increment of Δv2, and entered the outbound detailed inspection phase, conducting a close inspection of the target in the west and east directions along the light. The speed increment Δv2 determined the outbound approach distance.

[0058] Our satellite completed the third maneuver at position 3, with a velocity increment of Δv3, entered the U-turn phase, and changed the drift direction relative to the geosynchronous orbit belt;

[0059] Our satellite completed the fourth maneuver at position 4, with a speed increment of Δv4, and entered the return detailed inspection phase. The speed increment of Δv4 determined the return approach distance;

[0060] Our satellite completed the fifth maneuver at position 5 with a speed increment of Δv5. The speed increment Δv5 determines the return drift rate.

[0061] (2) Determine the time of each maneuver. The present invention determines the time of each approach according to the fixed position of the target to ensure observation in the direction of light.

[0062] Define the time when our satellite is at position 0 as T0, the orbital element as the semi-major axis a0, the eccentricity as e0, the inclination as i0, the right ascension of the ascending node as Ω0, the argument of perigee as ω0, the true anomaly as f0, and the geographical longitude of the sub-satellite point as L0. The above parameters are the planning input conditions;

[0063] Define the target sub-satellite point geographical longitude as L b , input conditions for planning;

[0064] The determination of the second maneuver time is divided into two steps: time determination and date determination.

[0065] The time is determined as 12:00 noon local time at the geographical longitude L0 of the target sub-satellite point;

[0066] The date is determined as follows: If the time T0 is earlier than 12:00 noon local time, then If the time of T0 is later than 12:00 noon when the star is below us, then Among them, R0=42164.5km is the semi-major axis of the geosynchronous orbit, and ceiling means the result is rounded up.

[0067] The first maneuver time is T1 = T2-24h;

[0068] The third maneuver time is T3 = T2 + n·24h, where n = 2, 3, 4, etc. The value of n needs to take into account both the concealment and timeliness of observation. The larger n is, the farther the turning position is from the target, the higher the concealment, but the observation interval between the outbound and return trips increases, and the timeliness of observation decreases.

[0069] The fourth maneuver time is T4 = T3 + 12h;

[0070] The fifth maneuver time is T5=T4+n·24h, where the value of n is the same as above.

[0071] (3) Optimizing and solving the first and second orbital maneuvering speed increments. The present invention determines the first two maneuvering speed increments based on the approach distance and the maneuvering time determined in (2).

[0072] Define our star to move to position 1. Before the first maneuver, the orbital element is the semi-major axis a 1_0 , eccentricity e 1_0 、Inclination angle i 1_0 , right ascension of ascending node Ω 1_0 、Arg of perigee ω 1_0 and true close angle f 1_0 ; The above parameters are solved by orbital numerical extrapolation method, where f 1_0 There is no analytical solution, so we can only solve it recursively. In the following formula, f is f0 and recursively reaches f 1_0 The intermediate value in the process, where μ = 398600.44 km 3 / s 2 is the gravitational constant at the center of the Earth:

[0073] a 1_0 =a0,e 1_0 =e0,i 1_0 =i0,Ω 1_0 =Ω0,ω 1_0 =ω0

[0074]

[0075] Among them, f 1_0 is the true anomaly angle of our star when it moves to position 1, μ is the gravitational constant of the Earth's center (μ=398600.44km 3 / s 2 ), Δt1 is the time length of the movement from the initial position to position 1, f is the true anomaly angle variable during the movement from the initial position to position 1, and t is the time variable during the movement from the initial position to position 1.

[0076] Define the maneuvering velocity increment of our satellite at position 1 as Δv1. Solve the initial position vector of our satellite at the first maneuvering position of position 1 in the Earth's inertial coordinate system and velocity vector And the velocity vector before the first maneuver Apply the first maneuver velocity increment Δv1 to obtain the velocity vector after the first maneuver The initial value of Δv1 is Δv 1_0 =0m / s.

[0077]

[0078]

[0079] In the Earth's inertial coordinate system, solve the orbital elements a1, e1, i1, Ω1, ω1, and f1 of our satellite after the first maneuver at the first maneuver position of position 1:

[0080] make

[0081]

[0082]

[0083] in, It is an intermediate calculation variable with no real physical meaning and does not need to be specially defined.

[0084] Solve the initial orbital elements of our satellite when it moves to the second maneuvering position at position 2:

[0085] a 2_0 =a1,e 2_0 =e1,i 2_0 =i1,Ω 2_0 =Ω1,ω 2_0 =ω1

[0086]

[0087] Define the maneuvering velocity increment of our satellite at position 2 as Δv2. Solve the initial position vector of our satellite at the second maneuvering position at position 2 in the Earth's inertial coordinate system. and velocity vector And apply the second maneuver speed increment Δv2 to obtain the speed vector after the second maneuver Where Δv2 is defined as a scalar, positive is the Tangent direction, negative is along The tangent direction is reversed.

[0088]

[0089]

[0090] According to the target star orbital element number a b (semi-major axis), e b (eccentricity), i b (inclination), Ω b (right ascension of ascending node), ω b (argument of perigee), f b (True Anomaly), solve the position vector of the target star in the Earth's inertial coordinate system

[0091]

[0092] Δv2 rough optimization:

[0093] The initial value of Δv2 is The outbound detailed inspection segment orbit crosses the geosynchronous orbit belt twice a day;

[0094] If the distance between the easternmost crossing point of the T2~T2+24h segment and the westernmost crossing point of the T2+24h~T2+48h segment is less than 2D0, continue to increase Δv2;

[0095] If the distance between the easternmost crossing point of the T2~T2+24h segment and the westernmost crossing point of the T2+24h~T2+48h segment is greater than 2D0, continue to reduce Δv2;

[0096] Until the spacing = 2D0, the rough optimal intermediate value Δv is obtained 2_1 ; Where D0 is the expected observation distance for detailed inspection of the target.

[0097] Δv1 and Δv2 are optimized:

[0098] Let t 下 =T2+24h, calculate t 下 Time and in, is the position vector of the Earth's center pointing to our star at T2+24h, It is the position vector pointing from the center of the earth to the target at T2+24h;

[0099] like exist On the east side, Δv1 is increased and Δv2 is reduced synchronously to ensure Δv1+Δv2=Δv 2_1 ;

[0100] like exist On the west side, Δv1 is reduced and Δv2 is increased synchronously to ensure Δv1+Δv2=Δv 2_1 ;

[0101] Until and In parallel, the final first two maneuver speed increments Δv1 and Δv2 are obtained.

[0102] (4) Determine the velocity increment Δv3 of the third orbital maneuver:

[0103]

[0104] The orbit extrapolation method after the third maneuver is similar to that in (3) and will not be repeated in the following text.

[0105] (5) Optimize and solve the fourth orbital maneuver velocity increment Δv4:

[0106] Take the initial value of Δv4 The return detailed inspection segment orbit crosses the geosynchronous orbit belt twice a day;

[0107] Let t 上 =T5-24h, calculate t 上 Time and Among them, t 上 is the time above the target after the fourth orbital maneuver, is the position vector of the Earth's center pointing to our star at T5-24h, It is the position vector of the center of the earth pointing to the target at T5-24h;

[0108] like exist On the east side, Δv4 continues to increase;

[0109] like exist On the west side, Δv4 continues to decrease;

[0110] Until and In parallel, the optimized result Δv4 is obtained.

[0111] (6) Determine the fifth orbital maneuver velocity increment Δv5:

[0112]

[0113] So far, the present invention has determined the time and speed increments of five orbital maneuvers based on the initial orbital elements of our satellite and the target, as well as the approach distance requirements, and completed the detailed investigation mission planning.

[0114] Example

[0115] The method for planning detailed investigation missions of geosynchronous orbit target imaging satellites proposed in the present invention utilizes a "water drop" shaped orbit with a large eccentricity. Under the premise of not conducting high-intensity maneuvers such as accompanying flight and circling flight, six-segment orbit splicing is used to complete four close-inspection observations of the target in high concealment, along the light, and in multiple directions.

[0116] The specific steps are:

[0117] In this example, the epoch of our satellite is 00:00:00 on June 4, 2023, with a semi-major axis of 42084km, eccentricity of 0, inclination of 0°, right ascension of ascending node of 0°, argument of perigee of 0°, and true anomaly of 144.896°. The target epoch is 00:00:00 on June 4, 2023, with a semi-major axis of 42164.5km, eccentricity of 0, inclination of 0°, right ascension of ascending node of 0°, argument of perigee of 0°, and true anomaly of 146.881°, and the fixed position is 75°W.

[0118] Determine the time of each maneuver. In this example, the second maneuver date of our satellite is determined to be June 5, 2023, and the time is 19:00:00. The first maneuver time is June 4, 2023, 19:00:00, the third maneuver time is June 8, 2023, 19:00:00, the fourth maneuver time is June 9, 2023, 7:00:00, and the fifth maneuver time is June 12, 2023, 7:00:00.

[0119] Determine the approach distance. In this example, the approach distance is selected as 30km.

[0120] The algorithm of the present invention is used to solve the problem, and finally the following is obtained:

[0121] The first maneuver speed increment is 0.42 m / s;

[0122] The second maneuver speed increment is 1.74 m / s;

[0123] The third maneuver speed increment was 0.78 m / s;

[0124] The speed increment for the fourth maneuver was 0.80 m / s;

[0125] The fifth maneuver speed increment was 2.10 m / s;

[0126] The total speed increment is 5.84m / s (while the speed increment consumption for accompanying and circling at a close distance of 30km is 9.32m / s, and the speed increment consumption in this example is optimized by 37%).

[0127] The first approach observation time is 1:55:00 on June 6, 2023, the approach distance is 29.7km, and the illumination angle is 22.7°;

[0128] The second approach observation time is 11:53:00 on June 7, 2023, the approach distance is 30.2km, and the illumination angle is 23.0°;

[0129] The third approach observation time is 13:36:00 on October 10, 2023, the approach distance is 30.4km, and the illumination angle is 23.1°;

[0130] The fourth approach observation time is 23:42:00 on October 11, 2023, the approach distance is 30.3km, and the illumination angle is 22.9°;

[0131] The observation concealment of the present invention is strong, and the satellite is always in a drifting state. Compared with the currently commonly used accompanying flight and circling flight strategies, there is no long-term approaching state with the target, which is conducive to concealing the observation intention, reducing the intensity of on-orbit game, avoiding legal disputes, and avoiding the target to be aware of and avoid observation as much as possible, which is conducive to achieving the purpose of observation; the present invention can implement 4 close-up detailed inspections from 4 different directions (west side facing the ground, west side facing away from the ground, east side facing the ground, and east side facing away from the ground) within a few days, and all are in the condition of direct light, and in addition, it can focus on observing the east and west sides of the target (conventional drift patrol can only observe a single target once, and focus on observing the target facing the ground and away from the ground, while the key payloads of most space situational awareness and space attack and defense satellites are located on the east and west sides), and the close-up reconnaissance process is dispersed within a few days, the target avoidance reconnaissance is more difficult and costly, and the observation result has a higher intelligence value; the present invention completes the observation of the same target under the same conditions, and compared with the currently commonly used accompanying flight and circling flight strategies, the speed increment consumption is reduced by more than 30%, which can greatly improve the reconnaissance efficiency of the whole satellite.

[0132] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A method for planning a detailed investigation mission of a geosynchronous orbit target imaging satellite, characterized in that include: The mission track is divided into an approaching section, a phasing section, an outbound detailed inspection section, a U-turn section, a return detailed inspection section, and a distance section; wherein the first orbital maneuver is performed between the approaching section and the phasing section, the second orbital maneuver is performed between the phasing section and the outbound detailed inspection section, the third orbital maneuver is performed between the outbound detailed inspection section and the U-turn section, the fourth orbital maneuver is performed between the U-turn section and the return detailed inspection section, and the fifth orbital maneuver is performed between the return detailed inspection section and the distance section; determining maneuver times for a first orbital maneuver, a second orbital maneuver, a third orbital maneuver, a fourth orbital maneuver, and a fifth orbital maneuver; Determine the velocity increments for the first orbital maneuver, the second orbital maneuver, the third orbital maneuver, the fourth orbital maneuver, and the fifth orbital maneuver.

2. The method for planning a geosynchronous orbit target imaging satellite focused detailed inspection mission according to claim 1, characterized in that: Determining the maneuvering time T2 of the second orbital maneuver includes determining the time and determining the date; wherein, The time is determined to be 12:00 noon local time at the geographical longitude L0 of the preset target sub-satellite point; The date to be determined includes: If the time T0 of our satellite at the initial position is earlier than 12:00 noon local time when our satellite is below the ground, then the maneuvering time of the second orbital maneuver is If the time of T0 is later than 12:00 noon when the star is below us, then Where R0 is the semi-major axis of the geosynchronous orbit, a0 is the semi-major axis, ceiling means the result is rounded up, L b is the geographical longitude of the target sub-satellite point.

3. The method for planning a geosynchronous orbit target imaging satellite focused detailed inspection mission according to claim 1, characterized in that: The first orbital maneuvering time is T1=T2-24h; wherein T1 is the first orbital maneuvering time, and T2 is the second orbital maneuvering time.

4. The method for planning a geosynchronous orbit target imaging satellite focused detailed inspection mission according to claim 1, characterized in that: The third orbital maneuvering time is T3=T2+n·24h; wherein T3 is the third orbital maneuvering time, and n is a positive integer greater than 1.

5. The method for planning a geosynchronous orbit target imaging satellite focused detailed inspection mission according to claim 1, characterized in that: The fourth orbital maneuvering time is T4=T3+12h; wherein T3 is the third orbital maneuvering time, and T4 is the fourth orbital maneuvering time.

6. The method for planning a geosynchronous orbit target imaging satellite focused detailed investigation mission according to claim 1, characterized in that: The fifth orbital maneuvering time is T5=T4+n·24h; wherein T5 is the fifth orbital maneuvering time, T4 is the fourth orbital maneuvering time, and n is a positive integer greater than 1.

7. The method for planning a geosynchronous orbit target imaging satellite focused detailed investigation mission according to claim 1, characterized in that: Determining the velocity increment for the first orbital maneuver and the velocity increment for the second orbital maneuver includes: Determine the initial value of the velocity increment Δv for the second orbital maneuver 2_0 , so that the outbound detailed inspection segment orbit crosses the geosynchronous orbit belt twice a day; The distance between the easternmost crossing point of the T2~T2+24h segment and the westernmost crossing point of the T2+24h~T2+48h segment is equal to 2D0, and the rough optimization intermediate value Δv of the velocity increment of the second orbital maneuver is obtained. 2_1 ; Where D0 is the expected observation distance for detailed inspection of the target; When T2+24h and In parallel, we obtain the velocity increment of the first orbital maneuver and the velocity increment of the second orbital maneuver; where, is the position vector of the Earth's center pointing to our star at T2+24h, It is the position vector pointing from the center of the earth to the target at T2+24h.

8. The method for planning a geosynchronous orbit target imaging satellite focused detailed investigation mission according to claim 1, characterized in that: The velocity increment Δv3 of the third orbital maneuver is obtained by the following formula: Where a0 is the semi-major axis, Δv1 is the velocity increment of the first orbital maneuver, Δv2 is the velocity increment of the second orbital maneuver, μ is the gravitational constant at the center of the Earth, and R0 is the semi-major axis of the geosynchronous orbit.

9. The method for planning a geosynchronous orbit target imaging satellite focused detailed investigation mission according to claim 1, characterized in that: The velocity increments for the fourth orbital maneuver are determined by: Determine the initial value of the velocity increment Δv for the fourth orbital maneuver 4_0 , so that the return detailed inspection segment orbit crosses the geosynchronous orbit belt twice a day; When T5-24h and In parallel, we get the velocity increment of the fourth orbital maneuver; where, is the position vector of the Earth's center pointing to our star at T5-24h, It is the position vector of the center of the earth pointing to the target at T5-24h.

10. The method for planning a geosynchronous orbit target imaging satellite focused detailed investigation mission according to claim 1, characterized in that: The velocity increment Δv5 of the fifth orbital maneuver is obtained by the following formula: Where a0 is the semi-major axis, Δv4 is the velocity increment of the second orbital maneuver, μ is the gravitational constant at the center of the Earth, and R0 is the semi-major axis of the geosynchronous orbit.

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