A patrol mission planning method for geosynchronous orbit target imaging satellite

Abstract

A method for planning a drift-and-survey mission for a geostationary orbit space target monitoring and imaging satellite is disclosed. The initial time and corresponding initial position of the observation satellite's roving mission are determined based on requirements, as well as the approach observation time and corresponding approach observation position. At the first maneuver time, the observation satellite moves to the first maneuver position, and a first velocity increment is applied to the initial velocity vector at this point, causing the observation satellite to maneuver tangentially along the orbit to obtain the velocity vector after the first maneuver. At the second maneuver time, the observation satellite moves to the second maneuver position, and a second velocity increment is applied to the initial velocity vector at this point, causing the observation satellite to maneuver tangentially along the orbit to obtain the velocity vector after the second maneuver. At the approach observation time, the observation satellite is at the approach observation position and observes the target satellite under front lighting conditions using the observation vector. This invention can improve roving efficiency and flexibility, realize drift-and-survey observation of targets on different surfaces, save propellant, and improve satellite performance.

Description

Technical Field

[0001] The present invention belongs to the technical field of space situation awareness space target monitoring and imaging, and in particular relates to a method for planning drift patrol missions of geosynchronous orbit space target monitoring and imaging satellites. Technical Background

[0002] Drift patrol is one of the most commonly used mission modes for geosynchronous orbit space target monitoring and imaging satellites. 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 monitoring and imaging of the target when drifting under the target. The satellite usually uses more than 100 days to complete drift patrols of multiple targets in the control area. It is a commonly used survey method. The satellite must be in a forward-light state during imaging, that is, the angle between the vector from the satellite to the target and the sunlight vector should be as small as possible to ensure the imaging quality. When the satellite is drifting and patrolling, it is necessary to adjust the orbit altitude specifically for the target. One purpose is to adjust the approach imaging distance. The second purpose is to ensure that the satellite is in a forward-light state when it moves to the approach imaging position by selecting the appropriate orbit altitude adjustment time.

[0003] At present, the drift patrol orbit altitude adjustment of geosynchronous orbit space target monitoring imaging satellite adopts the Holman orbit change method, which raises the satellite circular orbit altitude through two orbital maneuvers. This method has the following limitations: ① The propellant consumption is large and the number of maneuvers is large. A total of 4 maneuvers are required for approach and withdrawal; ② The concealment, timeliness and efficiency are poor. Due to the large number of maneuvers, it is usually necessary to start orbit altitude adjustment at least 1 day in advance. The intention is easy to be recognized by the opponent in advance and the timeliness is poor; ③ The mission flexibility is poor and the observation direction is single. The satellite is at the shortest distance when it passes directly under the target. Therefore, it can generally only observe the target's direction to the ground, and cannot observe other directions of the target, which cannot meet the flexible and diverse observation needs and intelligence needs; ④ The satellite needs to maintain a small orbital inclination in orbit to meet the observation needs of the target. The propellant demand for north-south position maintenance is large, which restricts the satellite's in-orbit maneuverability and patrol efficiency. Summary of the invention

[0004] The technical problem solved by the present invention is that a method for planning a drift patrol mission of a traditional space target monitoring and imaging satellite by adjusting the height of a circular orbit is proposed to address the prominent problems that restrict patrol effectiveness, such as large satellite propellant consumption, patrol concealment, poor timeliness and efficiency, and poor mission flexibility. A method for planning a satellite drift patrol mission based on adjusting the semi-major axis is proposed.

[0005] The technical solution of the present invention is:

[0006] A method for planning a drift patrol mission of a geosynchronous orbit space target monitoring imaging satellite is mainly to control the observation satellite to move to a close observation position through two maneuvering orbit changes, specifically:

[0007] First, according to the mission requirements, the initial time T0 and the corresponding initial position of the observation satellite patrol mission are determined, and the approach observation time T3 and the corresponding approach observation position are determined. The second maneuvering time T2 is determined by iterative optimization according to the approaching observation time T3, and then the first maneuvering time T1 is determined;

[0008] At the first maneuver time T1, the observation satellite moves to the first maneuver position The initial velocity vector here Applying the first velocity increment Δv1 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the first maneuver At the second maneuver time T2, the observation satellite moves to the second maneuver position The initial velocity vector here Applying the second velocity increment Δv2 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the second maneuver At the approach observation time T3, the observation satellite is in the approach observation position And the observation vector Observe the target satellite.

[0009] Furthermore, the approach observation time T3 and the corresponding approach observation position Determined by:

[0010] When the observation satellite and the target satellite are in the same orbital plane, looking from north to south in the Earth's inertial coordinate system, the Earth's rotation direction, the Sun's rotation direction around the Earth, and the satellite and template's rotation direction around the Earth are all counterclockwise. The moment when the observation satellite can observe along the light is taken as the approach observation moment T3;

[0011] When the observation satellite and the target satellite are not in the same plane, the time when the target passes through the observation satellite orbital plane is calculated and used as the approach observation time T3;

[0012] According to the approach observation time T3, the corresponding approach observation position is determined

[0013] Furthermore, it is determined that the observation satellite is in a close observation position. The observation vector of the target satellite include:

[0014] First, the position vector of the target satellite at time T3 is solved in the Earth's inertial coordinate system according to the orbital elements of the target satellite.

[0015]

[0016] Among them, ab is the semi-major axis, e b is the eccentricity, i b is the inclination angle, Ω b is the right ascension of the ascending node, ω b is the argument of perigee, f b is the true anomaly angle;

[0017] Then, determine the observation vector of the observation satellite to the target satellite at time T3

[0018]

[0019] in, It is the observation vector of the observation satellite to the target satellite at time T3 in the Earth's inertial coordinate system.

[0020] Furthermore, the determining of the first maneuvering time T1 specifically includes:

[0021] First, the optimization initial value of the first maneuver time T1 is specified as T 1_0 The initial optimization value of the second maneuver time T2 is T 2_0 ;

[0022] Select the initial value T 2_0 =T3-12h;

[0023] In T 2_0 Determine T based on 1_0 :If T 2_0 If the non-full-day portion of the time difference from T0 is greater than 12h, then T 1_0 =T 2_0 -n·24h+12h; if T 2_0 If the non-full-day portion of the time difference from T0 is less than or equal to 12 hours, then T 1_0 =T 2_0 -n·24h; n is a non-zero integer representing the whole day between two maneuvers. The value of n should ensure that T 1_0 The time difference with T0 is ≤24h.

[0024] Furthermore, the initial velocity vector Applying the first maneuver velocity increment Δv1 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the first maneuver Specifically include:

[0025] First, solve the initial position vector of the observation satellite moving to the first maneuver position in the Earth's inertial coordinate system: and the velocity vector

[0026]

[0027]

[0028]

[0029] The number of orbital elements before the observation satellite moves to the first maneuver position and performs the first maneuver: the semi-major axis is 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 ; t is the time variable.

[0030] Furthermore, the initial velocity vector Applying the second velocity increment Δv2 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the second maneuver Specifically include:

[0031] Solve the initial position vector of the observation satellite moving to the second maneuver position in the Earth's inertial coordinate system and the velocity vector

[0032]

[0033]

[0034]

[0035] Among them, Δv2 is a scalar, positive is the edge Tangent direction, negative is along The tangent direction is reversed; the initial value of Δv2 is 0m / s;

[0036] The initial orbital element of the observation satellite moving to the second maneuvering position: semi-major axis a 2_0 , eccentricity e 2_0 、Inclination angle i 2_0 , right ascension of ascending node Ω 2_0 、Arg of perigee ω 2_0 and true close angle f 2_0 ;

[0037] Then, T1 and T2 are further optimized and calculated according to the orbital period after the second maneuver:

[0038] ΔT is the orbital period after the second maneuver;

[0039] If the non-full-day portion of the time difference between T2 and T0 is less than or equal to 12 hours, then set T1 = T2-n·24 hours, where n is a non-zero integer representing the full-day portion between the two maneuvers, and the value of n is as large as possible to ensure that the time difference between T1 and T0 is less than or equal to 24 hours;

[0040] If the non-full-day portion of the time difference between T2 and T0 is greater than 12 hours, T1 is ordered to be T2-n·24h+12h, where n is a non-zero integer, and the value of n should ensure that the time difference between T1 and T0 is less than or equal to 24 hours.

[0041] Furthermore, the position vector of the observation satellite at the position close to the observation position It is expressed as:

[0042]

[0043] The orbital elements of the observation satellite moving to the close observation position are: semi-major axis a3, eccentricity e3, inclination i3, right ascension of ascending node Ω3, argument of perigee ω3 and true anomaly f3.

[0044] Furthermore, the first speed increment Δv1 and the second speed increment Δv2 are obtained by iterative optimization, including:

[0045] First, the first round of iterative optimization of the second maneuver velocity increment Δv2 is performed according to the orbital parameters after the second maneuver; including:

[0046] R E is the GEO orbit radius, is the observation vector of the observation satellite to the target satellite at time T3 in the Earth inertial coordinate system. The orbital elements of the observation satellite after the second maneuvering position are the semi-major axis a2 and the eccentricity e2.

[0047] like Then Δv2 increases in the positive direction, or decreases in the negative direction;

[0048] like Then increase Δv2 in the negative direction, or decrease Δv2 in the positive direction; return the observation vector of the observation satellite to the target satellite at time T3 Iterate the steps until Δv2 meets the optimization accuracy requirements;

[0049] Then, further optimizing Δv2 and optimizing Δv1 include:

[0050] Further optimize Δv2: when the calculated distance between the observation position and the center of the earth is smaller than the ideal distance between the observation position and the center of the earth, increase Δv2 in a positive direction or decrease Δv2 in a negative direction; when the calculated distance between the observation position and the center of the earth is larger than the ideal distance between the observation position and the center of the earth, increase Δv2 in a negative direction or decrease Δv2 in a positive direction; update T1, T2, Δt1, Δt2, Δt3 at the same time, and return the observation vector of the observation satellite to the target satellite at time T3 Solve again; where Δt1 = T1-T0, Δt2 = T2-T1, Δt3 = T3-T2;

[0051] Optimize Δv1: When the calculated observation position is east of the ideal observation position, increase Δv1 in a positive direction or decrease Δv1 in a negative direction; when the calculated observation position is west of the ideal observation position, increase Δv1 in a negative direction or decrease Δv1 in a positive direction; update T1, T2, Δt1, Δt2, and Δt3 at the same time, and return the observation vector of the observation satellite to the target satellite at time T3. Solve again; when

[0052] when and When the difference reaches the calculation accuracy requirement, the optimization solution is completed;

[0053] Finally, Δv2 and Δv1 are jointly optimized, including:

[0054] (a) Fix Δv1, optimize Δv2, and find the optimal Δv2 for the current Δv1;

[0055] (b) Fix Δv2, optimize Δv1, find the best Δv1 for the current Δv2, and then return to step (a) to loop and solve for the optimal solution;

[0056] when and When the difference reaches the calculation accuracy requirement, the optimization solution is completed.

[0057] The advantages of the present invention compared with the prior art are:

[0058] (1) The mission planning method for adjusting the semi-major axis of the orbit provided by the present invention only needs to adjust the satellite's quasi-synchronous circular orbit to an elliptical orbit, so that the satellite moves to the apogee closest to the target and has a straight-light condition. Only one orbit height adjustment maneuver is required. Compared with the existing method of raising the circular orbit through two maneuvers, the satellite speed increment consumption is reduced by about 50%, and the propellant consumption is reduced by about 50%.

[0059] (2) The task planning method for adjusting the semi-major axis of the orbit provided by the present invention requires only one orbit height adjustment maneuver, which is performed 12 hours before the close-in observation. Compared with the prior art, the shortened orbit height adjustment time can improve the concealment, timeliness and efficiency of the patrol.

[0060] (3) The task planning method for adjusting the semi-major axis of the orbit provided by the present invention can realize the observation of the target from any direction within the orbital plane, greatly improving the patrol flexibility and intelligence value.

[0061] (4) The mission planning method for adjusting the semi-major axis of the orbit provided by the present invention can realize drift patrol observation of non-planar targets, has no restrictions on the satellite's own inclination angle and the target's inclination angle, and has strong adaptability.

[0062] (5) The mission planning method for adjusting the semi-major axis of the orbit provided by the present invention allows the satellite to avoid north-south station maintenance while in orbit, and even to be pre-set at a certain inclination during the orbit entry phase. The propellant used for the satellite's north-south station maintenance and orbit change can be converted into propellant for the on-orbit patrol mission, greatly improving the satellite patrol efficiency. BRIEF DESCRIPTION OF THE DRAWINGS

[0063] Figure 1 The task planning method flow chart of the present invention;

[0064] Figure 2 The relative motion trajectory of the satellite in the target orbit coordinate system;

[0065] Figure 3 Schematic diagram of the relationship between the satellite observation target direction and the local time of the target sub-satellite point (looking north in the target satellite orbit coordinate system). DETAILED DESCRIPTION

[0066] The invention provides a method for planning a drift patrol mission of a geosynchronous orbit space target monitoring imaging satellite.

[0067] Our observation satellite completes the orbit change and drift action of the target satellite through two maneuvers. By reasonably selecting the timing of the maneuvers, the effects of the two maneuvers on the orbit are decoupled. The first maneuver is a phase adjustment, and the second maneuver is an orbit height adjustment, so that the observation satellite is in a close observation position when it moves to the apogee and has the conditions of direct sunlight.

[0068] First, according to the mission requirements, the initial time T0 and the corresponding initial position of the observation satellite patrol mission are determined, and the approach observation time T3 and the corresponding approach observation position are determined. The second maneuvering time T2 is determined by iterative optimization according to the approaching observation time T3, and then the first maneuvering time T1 is determined;

[0069] At the first maneuver time T1, the observation satellite moves to the first maneuver position The initial velocity vector here Applying the first velocity increment Δv1 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the first maneuver At the second maneuver time T2, the observation satellite moves to the second maneuver position The initial velocity vector here Applying the second velocity increment Δv2 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the second maneuver At the approach observation time T3, the observation satellite is in the approach observation position And the observation vector Observe the target satellite.

[0070] Execution steps are as follows Figure 1 As shown, specifically:

[0071] (1) Determine the number of maneuvers of our observation satellite to be 2, and define the four key positions of the observation satellite (i.e., initial position, first maneuver position, second maneuver position, and approach observation position) and their times, as shown in the attached figure. Figure 2 As shown:

[0072] The initial time of observing the satellite is T0, at which the observing satellite is at its initial position (position T0 in the figure);

[0073] The first maneuver time is T1, at which time the observation satellite is at the first maneuver position (position T1 in the figure), and performs the first maneuver along the orbital tangent;

[0074] The second maneuver time is T2, at which time the observation satellite is at the second maneuver position (position T2 in the figure), and performs the second maneuver along the orbital tangent direction;

[0075] Approach observation time T3, at this time the observation satellite is at the approach observation position (position T3 in the figure), which is the approach target time and is in the forward light state;

[0076] Let Δt1=T1-T0, Δt2=T2-T1, Δt3=T3-T2.

[0077] (2) Determine T0 and T3 times based on observation requirements. T0 time can be determined by the user based on requirements, T3 date depends on the timeliness requirements of the observation task, and the specific time of T3 depends on the orbits of the observation satellite and the target. For example, if the user wants to observe at 6:00, then this time is equivalent to the input of the optimization calculation. This observation time is related to the observation direction. The user is concerned about the observation direction, and the observation time can be determined based on the observation direction. The difference between T2 and T3 is half an orbital period. After T2 is determined, T1 is determined based on it. The selection principle depends on the non-whole-day part of the time difference between T2 and T0; then T1 is iteratively optimized.

[0078] The observation satellite and the target share the same orbital plane. At T3, the observation satellite is in a close observation position, and the observation time is related to the observation direction. Looking from north to south in the Earth's inertial coordinate system, the Earth's rotation direction, the Sun's rotation direction around the Earth, and the satellite and template's rotation direction around the Earth are all counterclockwise. Figure 3 As shown, in the target orbit coordinate system, looking from north to south: when the target satellite is at 0:00, the observation satellite is located directly below the target, and the observation can be achieved in the sunlight; when the target satellite is at 6:00, the observation satellite is located to the right of the target, and the observation can be achieved in the sunlight; when the target satellite is at 12:00, the observation satellite is located directly above the target, and the observation can be achieved in the sunlight; when the target satellite is at 18:00, the observation satellite is located to the left of the target, and the observation can be achieved in the sunlight. In fact, the user can observe the target satellite from any direction, and the observation time can be determined according to the observation direction as described above.

[0079] The observation satellite and the target satellite are not in the same plane. The time when the target passes through the observation satellite orbital plane is calculated as the approach time.

[0080] (3) According to the orbital element number a of the target satellite 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 angle), solve the position vector of the target satellite at time T3 in the Earth's inertial coordinate system

[0081]

[0082] (4) Based on And the observation time T3, the distance between the observation satellite and the target satellite, solve the observation position of the observation satellite at time T3 in is the observation vector of the satellite observing the target at time T3 in the Earth's inertial coordinate system:

[0083]

[0084] (5) Determine the initial values ​​of the iterative optimization of observation satellites T1 and T2:

[0085] T1 and T2 are the first and second maneuvering times of the observation satellite before approaching the target, which is also one of the optimization results of this method. The initial values ​​of the optimization are defined as T 1_0 and T 2_0 ;

[0086] Time T2 is defined as the last perigee before the observation satellite approaches the target. Then, at time T3, it reaches the apogee and observes the target satellite. The difference between time T2 and time T3 is half an orbital period, T2 = T3-12h. Since the observation satellite is in a quasi-synchronous orbit with a period slightly less than 24h, the initial value T is selected. 2_0 =T3-12h;

[0087] T 2_0 After determination, T is determined based on it. 1_0 , the selection principle depends on T 2_0 The non-full-day portion of the time difference from T0;

[0088] Such as T 2_0 If the non-full day portion of the time difference with T0 is less than or equal to 12 hours, then another T 1_0 =T 2_0 -n·24h, where n is a non-zero integer representing the entire day between two maneuvers. The value of n should be as large as possible to ensure that T 1_0 The time difference with T0 is ≤ 24h;

[0089] Such as T 2_0 If the non-full day portion of the time difference with T0 is greater than 12h, then another T 1_0 =T 2_0 -n·24h+12h

[0090] Where n is a non-zero integer representing the whole day between two maneuvers. The value of n should be as large as possible. Its value should ensure that T 1_0 The time difference with T0 is ≤24h.

[0091] (6) Before orbit extrapolation, the orbit parameters of the observation satellite are defined as follows:

[0092] The orbital elements at the initial position (T0) are: semi-major axis a0, eccentricity e0, inclination i0, right ascension of ascending node Ω0, argument of perigee ω0 and true anomaly f0;

[0093] Move to the first maneuver position (T1), before the first maneuver, the orbital root 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 ;

[0094] Move to the first maneuver position (T1). After the first maneuver, the orbital elements are the semi-major axis a1, eccentricity e1, inclination i1, right ascension of ascending node Ω1, argument of perigee ω1 and true anomaly f1;

[0095] Move to the second maneuver position (T2), before the second maneuver, the orbital root is the semi-major axis a 2_0 , eccentricity e 2_0 、Inclination angle i 2_0 , right ascension of ascending node Ω 2_0 、Arg of perigee ω 2_0 and true close angle f 2_0 ;

[0096] Move to the second maneuver position (T2). After the second maneuver, the orbital elements are semi-major axis a2, eccentricity e2, inclination i2, ascending node right ascension Ω2, perigee argument ω2 and true anomaly f2. The orbital period after the second maneuver is ΔT.

[0097] Move to the close observation position (T3), the orbital elements are semi-major axis a3, eccentricity e3, inclination i3, right ascension of ascending node Ω3, argument of perigee ω3 and true anomaly f3.

[0098] (7) Extrapolate to obtain the orbital elements before the first maneuver when the observation satellite moves to the first maneuver position, where f 1_0 There is no analytical solution, so it can only be solved by numerical integration. In the following formula, f is the recursive solution from f0 to f 1_0 The intermediate value in the process, only f 1_0 is related to the time variable t, and the other parameters are independent of time t:

[0099] a 1_0 =a0,e 1_0 =e0,i 1_0 =i0,Ω 1_0 =Ω0,ω 1_0 =ω0

[0100] Where μ is the gravitational constant at the center of the Earth.

[0101] (8) Solve the initial position vector of the observation satellite moving to the first maneuver position in the Earth's inertial coordinate system and the 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 Where Δv1 is defined as a scalar, and its value is Direction, negative is along The direction is reversed. The smaller the expected velocity increment, the better, which can save propellant, so the initial value of Δv1 is 0m / s:

[0102]

[0103]

[0104]

[0105] (9) Solve the orbital elements a1, e1, i1, Ω1, ω1, and f1 of the observation satellite after the first maneuver at the first maneuver position in the Earth inertial coordinate system:

[0106] Defining intermediate variables

[0107]

[0108]

[0109]

[0110]

[0111]

[0112]

[0113] (10) Extrapolate to solve the initial orbital elements when the observation satellite moves to the second maneuver position, only f 2_0 is related to time t, and the other parameters are independent of time t:

[0114] a 2_0 =a1,e 2_0 =e1,i 2_0 =i1,Ω 2_0 =Ω1,ω 2_0 =ω1

[0115] Here f is recursive from f1 to f 2_0 Intermediate values ​​in the process.

[0116] (11) Solve the initial position vector of the observation satellite moving to the second maneuver position in the Earth's inertial coordinate system and the velocity vector And applying the second maneuver speed increment Δv2, we get Where Δv2 is defined as a scalar, positive is the Tangent direction, negative is along The tangent direction is reversed. Similarly, the initial value of Δv2 can be 0m / s:

[0117]

[0118]

[0119]

[0120] (12) Solve the orbital elements of the observation satellite after the second maneuver at the second maneuver position in the Earth inertial coordinate system: semi-major axis a2, eccentricity e2, inclination i2, right ascension of ascending node Ω2, argument of perigee ω2 and true anomaly f2

[0121] Defining intermediate variables

[0122]

[0123]

[0124]

[0125]

[0126]

[0127]

[0128] (13) Recalculate T1 and T2 based on the actual orbit calculation results:

[0129]

[0130] If the non-full-day portion of the time difference between T2 and T0 is less than or equal to 12 hours, then set T1 = T2-n·24 hours, where n is a non-zero integer representing the full-day portion between two maneuvers of the observation satellite, and the value of n is as large as possible to ensure that the time difference between T1 and T0 is less than or equal to 24 hours;

[0131] If the non-full-day portion of the time difference between T2 and T0 is greater than 12 hours, T1 is ordered to be T2-n·24h+12h, where n is a non-zero integer representing the full-day portion between two maneuvers of the observation satellite, and the value of n is as large as possible to ensure that the time difference between T1 and T0 is less than or equal to 24 hours;

[0132] (14) Pre-optimization. Since the value of Δv2 has a great influence on the convergence of the algorithm, the first round of pre-iteration optimization of Δv2 is performed based on the orbital parameters after the second maneuver. E is the GEO orbit radius, is the observation vector of our satellite to the target at time T3 in the Earth's inertial coordinate system:

[0133] like Then it increases by Δv2 in a positive direction, or decreases by Δv2 in a negative direction (Δv2 is a scalar, which can be positive or negative. For example, a positive increase means that 1 changes to 1.5, and a negative decrease means that -1 changes to -0.8);

[0134] like Then Δv2 increases in the negative direction, or decreases in the positive direction;

[0135] Return to the above step (3) and iterate until Δv2 meets the optimization accuracy requirement.

[0136] (15) Extrapolate to solve the orbital elements when our satellite moves to the observation position. Only f3 is related to time t, and the other parameters are independent of time t:

[0137] a3=a2, e3=e2, i3=i2, Ω3=Ω2, ω3=ω2

[0138] Here f is the intermediate value in the process of recursively extending from f2 to f3.

[0139] (16) Solve the position vector of the observation satellite moving to the observation position in the Earth's inertial coordinate system

[0140]

[0141] (17) Rough optimization, and Compare and optimize Δv1 and Δv2 according to the comparison results:

[0142] Independently optimize Δv2. When the calculated observation position is lower than the ideal observation position (i.e., the calculated distance between the observation position and the center of the earth is smaller than the ideal distance between the observation position and the center of the earth), increase Δv2 in a positive direction or decrease Δv2 in a negative direction; when the calculated observation position is higher than the ideal observation position (i.e., the calculated distance between the observation position and the center of the earth is larger than the ideal distance between the observation position and the center of the earth), increase Δv2 in a negative direction or decrease Δv2 in a positive direction; simultaneously update T1, T2, Δt1, Δt2, Δt3, and return to step (3) to solve again;

[0143] Independently optimize Δv1. With the Earth's rotation direction from west to east as a reference, when the calculated observation position is eastward compared to the ideal observation position, increase Δv1 in a positive direction, or decrease Δv1 in a negative direction; when the calculated observation position is westward compared to the ideal observation position, increase Δv1 in a negative direction, or decrease Δv1 in a positive direction; simultaneously update T1, T2, Δt1, Δt2, Δt3, and return to step (3) to solve again;

[0144] and When the difference reaches the calculation accuracy requirement, the optimization solution is completed.

[0145] (18) Fine optimization: jointly optimize Δv2 and Δv1.

[0146] After completing the rapid rough optimization of Δv2 and Δv1 respectively, Δv2 and Δv1 are jointly optimized in two steps: (a) fix Δv1, optimize Δv2, and find the optimal Δv2 for the current Δv1; (b) fix Δv2, optimize Δv1, and find the optimal Δv1 for the current Δv2, and then return to (a) to loop and solve for optimization;

[0147] and When the difference reaches the calculation accuracy requirement, the optimization solution is completed.

[0148] Example

[0149] The specific steps of the monitoring imaging satellite drift patrol mission planning method of the present invention are:

[0150] (1) In this example, the target epoch is 19:52:00 on January 5, 2021, with a semi-major axis of 42164.47 km, an eccentricity of 0.000062, an inclination of 5.01°, a right ascension of the ascending node of 106.886°, an argument of perigee of 235.705°, and a mean anomaly of 126.260°. The epoch of our satellite is 19:52:00 on January 5, 2021, with a semi-major axis of 42081.58 km, an eccentricity of 0.000784, an inclination of 0.012°, a right ascension of the ascending node of 133.971°, an argument of perigee of 142.625°, and a mean anomaly of 188.901°.

[0151] (2) Determine the observation date. In this example, the approach date of our observation satellite is selected as January 8. At the same time, the observation satellite and the target are not in the same plane, so the time when the observed target crosses the observation satellite orbital plane is calculated as the time when the target is observed. In this example, the target orbit and the observation satellite orbital plane intersect twice between 0:00 and 24:00 on January 8, and one of them, 19:32:14, is selected as the observation time T3.

[0152] (3) Determine the approach distance. In this example, the approach distance is selected as 50 km.

[0153] (4) Determine the initial values ​​of the iterative optimization of the observation satellites T1 and T2. In this example, the initial value of the first maneuver time T1 is 7:32:14 on January 6, 2021, and the initial value of the second maneuver time T2 is 7:32:14 on January 8, 2021.

[0154] (5) Using the method of the present invention to solve, we finally get:

[0155] The first maneuver time is: 8:47:14 on January 6, 2021, speed increment -0.265m / s;

[0156] The second maneuver time is: 8:47:14 on January 8, 2021, with a speed increment of 0.315m / s;

[0157] The approach distance is 50.042km;

[0158] The illumination angle at the time of observation was 1.8°.

[0159] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A method for planning a drift patrol mission of a geosynchronous orbit space target monitoring imaging satellite, characterized in that: Control the observation satellite to make it move to the observation position through two maneuvers. First, according to the mission requirements, the initial time T0 and the corresponding initial position of the observation satellite patrol mission are determined, and the approach observation time T3 and the corresponding approach observation position are determined. The second maneuvering time T2 is determined by iterative optimization according to the approaching observation time T3, and then the first maneuvering time T1 is determined; At the first maneuver time T1, the observation satellite moves to the first maneuver position The initial velocity vector here Applying the first velocity increment Δv1 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the first maneuver At the second maneuver time T2, the observation satellite moves to the second maneuver position The initial velocity vector here Applying the second velocity increment Δv2 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the second maneuver At the approach observation time T3, the observation satellite is in the approach observation position And the observation vector Observe the target satellite; The approach observation time T3 and the corresponding approach observation position Determined by: When the observation satellite and the target satellite are in the same orbital plane, looking from north to south in the Earth's inertial coordinate system, the Earth's rotation direction, the Sun's rotation direction around the Earth, and the satellite and template's rotation direction around the Earth are all counterclockwise. The moment when the observation satellite can observe along the light is taken as the approach observation moment T3; When the observation satellite and the target satellite are not in the same plane, the time when the target passes through the observation satellite orbital plane is calculated and used as the approach observation time T3; According to the approach observation time T3, the corresponding approach observation position is determined Confirm that the observation satellite is in a close observation position The observation vector of the target satellite include: First, the position vector of the target satellite at time T3 is solved in the Earth's inertial coordinate system according to the orbital elements of the target satellite. Among them, a b is the semi-major axis, e b is the eccentricity, i b is the inclination angle, Ω b is the right ascension of the ascending node, ω b is the argument of perigee, f b is the true anomaly angle; Then, determine the observation vector of the observation satellite to the target satellite at time T3 in, It is the observation vector of the observation satellite to the target satellite at time T3 in the Earth's inertial coordinate system.

2. The method for planning a geosynchronous orbit space target monitoring imaging satellite drift patrol mission according to claim 1, characterized in that: Determining the first maneuvering time T1 specifically includes: First, the optimization initial value of the first maneuver time T1 is specified as T 1_0 The initial optimization value of the second maneuver time T2 is T 2_0 ; Select the initial value T 2_0 =T3-12h; In T 2_0 Determine T based on 1_0 :If T 2_0 If the non-full-day portion of the time difference from T0 is greater than 12h, then T 1_0 =T 2_0 -n24h+12h; if T 2_0 If the non-full-day portion of the time difference from T0 is less than or equal to 12 hours, then T 1_0 =T 2_0 -n·24h; n is a non-zero integer representing the whole day between two maneuvers. The value of n should ensure that T 1_0 The time difference with T0 is ≤24h.

3. The method for planning a geosynchronous orbit space target monitoring imaging satellite drift patrol mission according to claim 2, characterized in that: The initial velocity vector Applying the first maneuver velocity increment Δv1 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the first maneuver Specifically include: First, solve the initial position vector of the observation satellite moving to the first maneuver position in the Earth's inertial coordinate system: and the velocity vector The number of orbital elements before the observation satellite moves to the first maneuver position and performs the first maneuver: the semi-major axis is 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 ; t is the time variable.

4. The method for planning a geosynchronous orbit space target monitoring imaging satellite drift patrol mission according to claim 3, characterized in that: The initial velocity vector Applying the second velocity increment Δv2 causes the observation satellite to maneuver along the orbital tangent direction to obtain the velocity vector after the second maneuver Specifically include: Solve the initial position vector of the observation satellite moving to the second maneuver position in the Earth's inertial coordinate system and velocity vector Among them, Δv2 is a scalar, positive is the edge Tangent direction, negative is along The tangent direction is reversed; the initial value of Δv2 is 0m / s; The initial orbital element of the observation satellite moving to the second maneuvering position: semi-major axis a 2_0 , eccentricity e 2_0 、Inclination angle i 2_0 , right ascension of ascending node Ω 2_0 、Arg of perigee ω 2_0 and true close angle f 2_0 ; Then, T1 and T2 are further optimized and calculated according to the orbital period after the second maneuver: ΔT is the orbital period after the second maneuver; If the non-full-day portion of the time difference between T2 and T0 is less than or equal to 12 hours, then set T1 = T2-n24h, where n is a non-zero integer representing the full-day portion between the two maneuvers, and the value of n is as large as possible to ensure that the time difference between T1 and T0 is less than or equal to 24 hours; If the non-full-day portion of the time difference between T2 and T0 is greater than 12 hours, T1 is ordered to be T2-n24h+12h, where n is a non-zero integer, and the value of n should ensure that the time difference between T1 and T0 is less than or equal to 24 hours.

5. The method for planning a geosynchronous orbit space target monitoring imaging satellite drift patrol mission according to claim 4, characterized in that: The position vector of the observation satellite at the position close to the observation position It is expressed as: The orbital elements of the observation satellite moving to the close observation position are: semi-major axis a3, eccentricity e3, inclination i3, right ascension of ascending node Ω3, argument of perigee ω3 and true anomaly f3.

6. The method for planning a geosynchronous orbit space target monitoring imaging satellite drift patrol mission according to claim 5, characterized in that: The first speed increment Δv1 and the second speed increment Δv2 are obtained by iterative optimization, including: First, the first round of iterative optimization of the second maneuver velocity increment Δv2 is performed according to the orbital parameters after the second maneuver; including: R E is the GEO orbit radius, is the observation vector of the observation satellite to the target satellite at time T3 in the Earth inertial coordinate system. The orbital elements of the observation satellite after the second maneuvering position are the semi-major axis a2 and the eccentricity e2. like Then Δv2 increases in the positive direction, or decreases in the negative direction; like Then increase Δv2 in the negative direction, or decrease Δv2 in the positive direction; return the observation vector of the observation satellite to the target satellite at time T3 Iterate the steps until Δv2 meets the optimization accuracy requirement; Then, further optimizing Δv2 and optimizing Δv1 include: Further optimize Δv2: when the calculated distance between the observation position and the center of the earth is smaller than the ideal distance between the observation position and the center of the earth, increase Δv2 in a positive direction or decrease Δv2 in a negative direction; when the calculated distance between the observation position and the center of the earth is larger than the ideal distance between the observation position and the center of the earth, increase Δv2 in a negative direction or decrease Δv2 in a positive direction; update T1, T2, Δt1, Δt2, Δt3 at the same time, and return the observation vector of the observation satellite to the target satellite at time T3 Solve again; where Δt1 = T1-T0, Δt2 = T2-T1, Δt3 = T3-T2; Optimize Δv1: When the calculated observation position is east of the ideal observation position, increase Δv1 in a positive direction or decrease Δv1 in a negative direction; when the calculated observation position is west of the ideal observation position, increase Δv1 in a negative direction or decrease Δv1 in a positive direction; update T1, T2, Δt1, Δt2, and Δt3 at the same time, and return the observation vector of the observation satellite to the target satellite at time T3. Solve again; when when and When the difference reaches the calculation accuracy requirement, the optimization solution is completed; Finally, Δv2 and Δv1 are jointly optimized, including: (a) Fix Δv1, optimize Δv2, and find the optimal Δv2 for the current Δv1; (b) Fix Δv2, optimize Δv1, find the best Δv1 for the current Δv2, and then return to step (a) to loop and solve for the optimal solution; when and When the difference reaches the calculation accuracy requirement, the optimization solution is completed.

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