A non-cooperative target-oriented anti-collision hovering trajectory control method and system

Abstract

The present invention discloses a collision-avoiding hovering orbit control method for non-cooperative targets, including: step one, establishing a relative motion equation based on orbit root difference and a relative motion control equation based on Gaussian perturbation; step two, the mission satellite forms an anti-collision configuration through three pulse maneuvers to achieve rapid approach from long distance to short distance; step three, the mission satellite performs eccentricity error correction after approaching from a long distance; step four, the mission satellite further performs semi-major axis and eccentricity adjustment to prepare for hovering; step five, the mission satellite approaches to a designated hovering position through three pulse maneuvers. The method of the present invention can solve the problem that the prior art does not consider collision-avoiding approach control for non-cooperative target observation. The orbit control scheme is simple and reliable, easy to achieve high-precision control, and effectively avoids the risk of orbit control failure due to delays in orbit measurement and control and instability of data transmission.

Description

Technical Field

[0001] The invention relates to the technical field of spacecraft formation flying, and relates to an anti-collision hovering orbit control method and system. Background Art

[0002] With the development of aerospace technology and the increase in mission requirements, the research on tracking observation, space attack and defense, rendezvous and docking based on relative motion models has received more and more attention and importance. Realizing long-term, autonomous and precise guidance and control for non-cooperative maneuvering targets is conducive to space operations and mission completion, and is effectively applied to many space technology fields such as on-orbit services and satellite formations. The multi-pulse guidance and control of non-cooperative targets in space mainly studies the orbital motion between two close-range spacecraft, establishes the relative dynamics model between the mission satellite and the target satellite, and provides a multi-pulse guidance and control method under ground guidance.

[0003] Most application satellites are operating in near-circular orbits. In practice, due to various errors, the orbit can be approximated as a circular orbit and described by the average orbital elements. Orbital control uses thrust to make the orbital parameters produce the required changes. Orbital control for hovering observation of non-cooperative targets is mainly achieved by relative motion state. Existing methods usually use the observation information provided by the on-board optical camera to describe the observation maneuver control problem of non-cooperative targets as a standard nonlinear programming problem, optimize the approach path of the mission satellite, and realize hovering observation of the specified configuration of the mission satellite. Among them, the path optimization process relies on the on-board computer for real-time solution, which is a test for the performance, power consumption and reliability of the mission satellite. In addition, in the process of approaching the non-cooperative target, there is a risk of orbit control failure due to the delay of orbit measurement and control and the instability of data transmission. No effective solution has been proposed for the anti-collision problem during the rapid approach of the mission satellite. Summary of the invention

[0004] The technical problem solved by the present invention is: in view of the fact that the prior art does not consider the needs of anti-collision approach and hovering orbit control for non-cooperative targets, the present invention provides an anti-collision hovering orbit control method and system for non-cooperative targets. The method is simple and reliable, meets the actual needs of the project, and can meet the flexibility and rapidity requirements of the task in most cases.

[0005] The technical solution adopted by the present invention is: a collision-avoiding hovering track control method for a non-cooperative target, comprising the following steps:

[0006] Step 1: Establish the relative motion equation based on orbital root difference and the relative motion control equation based on Gaussian perturbation;

[0007] First, the orbital element difference of the spacecraft is converted into the relative position and relative velocity. The position of the mission satellite in the orbital reference coordinate system can be expressed as (x, y, z), and the orbital element of the mission satellite is defined as The orbital elements of the target star are The subscripts d and c represent the mission star and the target star respectively. The difference in the orbital elements between the two, that is, the relative orbital elements, is:

[0008] a d is the semi-major axis of the mission star, e d is the eccentricity of the mission star, i d is the orbital inclination of the mission star, Ω d is the right ascension of the mission star's ascending node, ω d is the mission star's argument of perigee, u d is the latitude argument of the mission satellite;

[0009] a c is the semi-major axis of the target star, e c is the eccentricity of the target star, i c is the orbital inclination of the target star, Ω c is the right ascension of the ascending node of the target star, ω c is the target star's perigee argument, u c is the latitude argument of the target star.

[0010] Δa, Δe, Δi, ΔΩ, Δω, and Δu are the relative semi-major axis, relative eccentricity, relative orbital inclination, relative right ascension of ascending node, relative argument of perigee, and relative argument of latitude, respectively;

[0011] Then the relative motion equation based on the orbital root difference is:

[0012]

[0013] Among them, u0 is the latitude argument of the mission satellite’s initial position, and the relative eccentricity vector and relative inclination vector are:

[0014]

[0015] For the mission satellite orbit control in a near-circular orbit, the change in the number of mission satellite orbit elements is δλ=[δa,δi,δΩ,δe x ,δe y ], then the relative motion control equation based on Gaussian perturbation is:

[0016]

[0017]

[0018]

[0019]

[0020]

[0021] Where V is the orbital velocity of the mission satellite, the subscripts R, T, and W are the radial, track, and normal directions, respectively, and ΔV R ,ΔV T ,ΔV W are the velocity increments in the three directions. Radial thrust is generally not used in actual orbit control. δa,δi,δΩ,δe x ,δe y They are respectively the change of semi-major axis, the change of orbital inclination, the change of right ascension of ascending node, the change of eccentricity in the x-axis direction, and the change of eccentricity in the y-axis direction;

[0022] Step 2: Control the mission satellite to form an anti-collision configuration through three pulse maneuvers, and quickly approach the target satellite from the long distance D0 to the short distance D1;

[0023] The mission star and the target star are orbiting in the same orbit, with the semi-major axis a d =a c =a, the initial position is at D0 (D0>0) behind the target satellite, and the mission satellite is designed to quickly approach the target satellite from D0 to D1 (D0>D1) behind the target satellite. In addition to the two pulses at the start and end points, in order to prevent the mission satellite from colliding with the target satellite during the approach process, a pulse configuration correction is required during the approach process. The specific steps are as follows:

[0024] S021, determine the transfer time of the mission satellite to the target satellite. Let the target satellite be D0 in n full orbit periods (T d ) approaches the target star D1, and the second maneuver is carried out n1 orbital periods after the first maneuver, where n1 is an odd multiple of 0.5.

[0025] S022, calculate the second maneuvering track pulse Δv T2 Let the x-direction safe distance between the mission star and the target star when the target star's latitude argument is 0 be Δx r , then considering the second maneuver trajectory pulse Δv under the atmospheric drag perturbation T2 Size:

[0026]

[0027] Among them, considering the atmospheric drag perturbation, the semi-major axis attenuation is estimated to obtain the pulse Δv required to offset the atmospheric drag at2 for

[0028]

[0029] Where ρ is the atmospheric density, C D is the drag coefficient of the spacecraft, ε is the surface-to-mass ratio, and the subscripts c and d represent the target satellite and the mission satellite, respectively.

[0030] S023, calculate the first maneuvering track pulse Δv T1 . Δv T1 The calculation formula is:

[0031]

[0032] Where, ΔL T1 The pulse Δv is taken into account to offset the atmospheric resistance. at1 The resulting displacement, ΔL T2 is the track displacement generated by the second maneuver, and the calculation formula is as follows

[0033] ΔL T1 =-3nT d Δv at1 , (7)

[0034]

[0035] ΔL T2 =-3(n-n1)T d Δv T2 (9)

[0036] S024, calculate the normal pulse Δv of the first maneuver W Let the z-direction safe distance between the mission star and the target star be Δz r , then the normal pulse of the first maneuver Δv W for

[0037]

[0038] S025, calculate the third maneuver track pulse quantity Δv T3 . Δv T3 The calculation formula is

[0039] Δv T3 =-(Δv T1 +Δv T2 ). (11)

[0040] S026, according to the calculation results of the above steps, apply the first maneuvering pulse (including the tracking pulse Δv T1 and the normal pulse Δv W ), in n1T d Apply the second maneuvering pulse Δv at time T2 , arrive near the target star D1 nTd Apply the third maneuvering pulse Δv at time T3 . Fly around D1 to turn on the onboard laser rangefinder and perform error correction.

[0041] Step 3: Control the mission satellite to perform eccentricity error correction at D1 after approaching from a long distance;

[0042] Due to the high-order perturbations of the Earth and the space environment, there is a large eccentricity error during the close approach process, and further eccentricity correction is required when flying around at D1. The specific steps for eccentricity error correction are:

[0043] S031, measurement mission satellite in u d = 0°, the components of the eccentricity error on the x and y axes are δe x ,δe y ;

[0044] S032, calculate the components of the total pulse on the x and y axes required for eccentricity error correction, which are Δv Tex , Δv Tey :

[0045]

[0046]

[0047] S033, in u d = 0°, apply correction pulse If this is the first correction, combine this correction pulse with the third pulse mentioned in step 2 otherwise,

[0048] S034, in u d =90°, apply correction pulse

[0049] S035, in u d =180°, apply correction pulse

[0050] S036, in u d =270°, apply correction pulse

[0051] S037, measurement mission star u d = 0° eccentricity error, if it is not zero, execute step S032, otherwise, execute step S038;

[0052] S038, after completing the correction, the mission star is in u d= 0° when applying correction pulse Δv at3 To offset atmospheric perturbations, the aircraft continued to fly around D1 for one cycle before making preparations for a close approach.

[0053] Δv at3 The calculation formula is:

[0054]

[0055] Where n e To correct the total period of the flyby at D1.

[0056] Step 4: Control the mission satellite to further adjust the semi-major axis and eccentricity at D1 to prepare for hovering;

[0057] Set the hovering position to be L behind the target star B , left side L L , below L D First, we need to eliminate the anti-collision safety configuration in the x and z directions at the same time, respectively at u d =0°,u d =180°The fourth and fifth orbit control pulse maneuvers are carried out at two moments. The fourth maneuver has a pulse quantity Δv T4 and the fifth maneuver trace pulse Δv T5 The calculation formula is as follows

[0058]

[0059]

[0060] Where, ΔL T5 is the track displacement produced by the fifth maneuver,

[0061]

[0062] ΔL T4 The track displacement caused by the advance offset of the semi-major axis in the fourth maneuver, in is the actual distance between the mission star and the target star during the fourth maneuver. After adjusting the eccentricity, the mission star is d =0° applied pulse-(Δv T4 +Δv T5 ), so that the mission star's orbital altitude remains unchanged, and this pulse is merged into the mission star's sixth maneuvering track pulse Δv T6 .

[0063] Step 5: Control the mission satellite to approach the designated hovering position from D1 through a three-pulse maneuver.

[0064] According to the law of relative motion, the mission star always has a drift in the velocity direction relative to the target star, and cannot be stably maintained at any relative position, but this property can be used to make the mission star drift to the specified position. The specific steps are as follows:

[0065] S051, calculate the required tracking pulse Δv during hovering D The calculation formula is as follows

[0066]

[0067] S052, calculate the pulse size Δv required for phase adjustment p . Δv D The resulting track drift ΔL D for

[0068]

[0069] Note that the semi-major axis offset displacement ΔL produced by offsetting atmospheric drag is ats , the phase adjustment needs to be compensated and the distance is configured as

[0070] ΔL s =ΔL D +L B +ΔL ats , (18)

[0071] Therefore, the pulse size required for phase adjustment is calculated as

[0072]

[0073] Where η is the mission star angular velocity.

[0074] S053, calculate the left bias normal pulse Δv Wp , the calculation formula is:

[0075]

[0076] S054, calculate the mission satellite's sixth maneuver track pulse Δv T6 , the seventh maneuver trace pulse Δv T7 and the normal pulse Δv W7 , the eighth maneuver trace pulse Δv T8 The calculation formula is:

[0077] Δv T6 =Δv p -Δv T4 -Δv T5

[0078] Δv T7 =-Δv p+Δv D ,Δv W7 =Δv Wp

[0079] Δv T8 =Δv D . (twenty one)

[0080] S055, calculate the results according to the above steps, and then d = 0°, the sixth motor tracking pulse Δv is applied T6 , after one cycle, the seventh maneuvering track and normal pulse Δv is applied T7 ,Δv W7 , on the mission star u d =180°, the eighth motor tracking pulse Δv is applied T8 , then when u d =270°, the mission star can reach the designated position on the left.

[0081] The control system according to the above-mentioned anti-collision hovering track control method for non-cooperative targets includes:

[0082] The first module is used to establish the relative motion equation based on orbital root difference and the relative motion control equation based on Gaussian perturbation;

[0083] The second module is used to control the mission satellite to form an anti-collision configuration through pulse maneuvers, so that the mission satellite can approach the target satellite from a long distance to a short distance.

[0084] The third module is used to control the mission satellite to perform eccentricity error correction after approaching from a long distance;

[0085] The fourth module is used to control the mission satellite to adjust the semi-major axis and eccentricity through pulse maneuvers;

[0086] The fifth module is used to control the mission satellite to approach the designated hovering position through pulse maneuvers.

[0087] The advantages of the present invention compared with the prior art are:

[0088] (1) The present invention proposes a multi-pulse guidance control method based on the Gaussian perturbation equation, taking the average orbital element as the controlled variable, while considering the high-order perturbation force of the earth and the perturbation force of the space environment. The algorithm is simple and reliable, and it is easy to achieve high-precision control;

[0089] (2) The present invention is aimed at the needs of hovering observation of non-cooperative targets, fully considering the limitations of the actual mission satellite's maneuverability and ground orbit determination capabilities, and designs a multi-pulse approach hovering orbit control method in detail. This pulse control method can simply and reliably achieve relative orbit control under ground guidance;

[0090] (3) The present invention takes into account the anti-collision requirements of the mission satellite during the approach process, and forms a safe approach configuration facing the non-cooperative target through track and normal pulses, effectively avoiding the risk of orbit control failure due to delays in orbit measurement and control and instability in data transmission. BRIEF DESCRIPTION OF THE DRAWINGS

[0091] Figure 1 is a flow chart of the track control method;

[0092] Figure 2 This is the relative trajectory diagram of the mission satellite's rapid approach from 50km to 4km;

[0093] Figure 3 is the anti-collision safety configuration (xz plane) diagram of the mission satellite during its rapid approach;

[0094] Figure 4 This is the mission satellite's 4km correction and relative orbital trajectory diagram;

[0095] Figure 5 This is the relative trajectory diagram of the mission satellite approaching the designated position at 4km and hovering for observation. DETAILED DESCRIPTION

[0096] The present invention will be further described in detail below with reference to the accompanying drawings.

[0097] Aiming at the requirements of approach, flyby and hovering observation in the hovering observation mission profile of non-cooperative targets, the present invention fully considers the limitations of the actual mission satellite's maneuverability and ground orbit determination capability, and designs a collision-avoiding hovering orbit control method in detail. The pulse control method can simply and effectively realize relative orbit control under ground guidance, which provides a certain reference and reference for the relative orbit design and control of space non-cooperative targets.

[0098] The present invention provides a non-cooperative target-oriented anti-collision hovering trajectory control method, such as Figure 1 A flow chart of the track control method is given, which includes the following steps:

[0099] Step 1: Establish the relative motion equation based on orbital root difference and the relative motion control equation based on Gaussian perturbation;

[0100] The on-orbit state of a spacecraft can be described by orbital elements. Similarly, the relative state of the mission satellite and the target satellite can also be described by the orbital element difference at each moment. Due to the limitation of existing measurement methods, the orbital elements cannot be directly observed. First, the orbital element difference is converted into relative position and relative velocity. When relative position and relative velocity are used to describe the relative motion law between spacecraft, the origin of the orbital reference coordinate system is defined at the center of mass of the target satellite, the x-axis is along the radial direction, the z-axis is along the direction of the orbital motion angular momentum, and the y-axis forms a right-handed orthogonal coordinate system with the x-axis and the z-axis. The position of the mission satellite in the reference coordinate system can be expressed as (x, y, z).

[0101] The orbital elements of the mission star are defined as The orbital elements of the target star are The subscripts d and c represent the mission star and the target star respectively. The difference in the orbital elements between the two, that is, the relative orbital elements, is:

[0102] a d is the semi-major axis of the mission star, e d is the eccentricity of the mission star, i d is the orbital inclination of the mission star, Ω d is the right ascension of the mission star's ascending node, ω d is the mission star's argument of perigee, u d is the latitude argument of the mission satellite;

[0103] a c is the semi-major axis of the target star, e c is the eccentricity of the target star, i c is the orbital inclination of the target star, Ω c is the right ascension of the ascending node of the target star, ω c is the target star's perigee argument, u c is the latitude argument of the target star.

[0104] The relative eccentricity vector and relative inclination vector are:

[0105]

[0106] Δe x , Δe y are the relative eccentricity vectors Components in the x and y axis directions; Δi x ,Δi y Relative inclination vector Components in the x and y axis directions;

[0107] Then the relative position, relative velocity and orbital root number difference have the following relationship:

[0108]

[0109] Where u0 is the latitude argument of the initial position of the mission satellite. When the semi-major axes of the orbits of the target satellite and the mission satellite are equal, the relative motion equation can be simplified to:

[0110]

[0111]

[0112] z=Bsin(ηt-θ) (24)

[0113] Where η is the mission satellite angular velocity, t is time, θ=arctan(Δi y ,Δi x ).

[0114] For the mission satellite orbit control in a near-circular orbit, the change in the mission satellite orbit element is [δa,δi,δΩ,δe x ,δe y ], the Gaussian perturbation equation can be used to describe the approximate dynamic model:

[0115]

[0116]

[0117]

[0118]

[0119]

[0120] Where V is the orbital velocity of the mission satellite, R, T, and W are the radial, track, and normal directions, respectively, and ΔV R ,ΔV T ,ΔV W is the velocity increment in three directions. If only the thrust in the direction of the trajectory is used, the following will be affected:

[0121]

[0122]

[0123]

[0124] If only normal thrust is used:

[0125]

[0126]

[0127] The two track pulse velocity increments can simultaneously adjust the three parameters of semi-major axis, eccentricity, and perigee angle, and the reasonable arrangement of the positions and sizes of the two pulses can also minimize the total velocity increment. Applying control in the normal direction can achieve control of the lateral motion amplitude. It is worth mentioning that radial thrust is generally not used in actual orbit control.

[0128] Step 2: The mission satellite forms an anti-collision configuration through three pulse maneuvers and achieves a rapid approach from the long distance D0 to the short distance D1;

[0129] The mission star and the target star are orbiting in the same orbit, with the semi-major axis a d =a c =a, the initial position is at D0 (D0>0) behind the target star, the orbit latitude argument is 0°, and the mission star is designed to quickly approach from D0 to D1 (D0>D1) behind the target star. In order to prevent the mission star from colliding with the target star during the approach process, it should be ensured that the mission star always maintains a certain distance from the target star in the xz plane, that is, the relative orbit configuration in the xz plane should be an elliptical anti-collision safety configuration, rather than a straight line. Therefore, the mission star needs to perform two maneuvers from D0 to D1 respectively. At time t0, the first maneuver pulse mainly provides a pulse for approaching the target star in the track direction, and the second maneuver further corrects the configuration to ensure the distance between the mission star and the target star in the xz plane.

[0130] Let the target star be D0 in n complete orbital periods (T d ) Approach the target star D1, and the second maneuver is performed n1 orbital periods after the first maneuver, where n1 is an odd multiple of 0.5. The anti-collision safety configuration of the approach is mainly achieved through the second maneuver. The distance between the mission star and the target star has a certain safety distance in the xz plane, and the orbital projection is an ellipse. First, determine the pulse amount applied by the second maneuver. The specific steps are as follows: Let the x-direction safety distance between the mission star and the target star when the latitude angle of the target star is 0 be Δx r V is the orbital velocity of the mission star, and the relationship between the x direction and the pulse is

[0131]

[0132] Calculate the second ideal maneuver pulse

[0133]

[0134] Considering the atmospheric drag perturbation, the semi-major axis attenuation is estimated to obtain the pulse required to offset the atmospheric drag:

[0135]

[0136] Where ρ is the atmospheric density, C D is the drag coefficient of the spacecraft, ε is the surface-to-mass ratio, and the subscripts c and d represent the target satellite and the mission satellite respectively. Then the second maneuver pulse Δv T2 :

[0137]

[0138] The first maneuver mainly provides the speed of the target satellite approaching the mission satellite, and the speed of the process from D0 to D1 is obtained by summing the track pulses of the first and second maneuvers. The track displacement generated by the second maneuver is:

[0139] ΔL T2 =-3(n-n1)T d Δv T2 , (32)

[0140] Consider the pulse Δv to offset the atmospheric drag at1 ,

[0141]

[0142] The resulting displacement is obtained:

[0143] ΔL T1 =-3nT d Δv at1 (34)

[0144] Therefore, the first maneuver pulse Δv is calculated T1 :

[0145]

[0146] At the same time, since the motion in the z direction is decoupled from the x direction, a z-direction pulse is applied during the first maneuver to ensure the safe distance in this direction. In engineering, the z-direction adjustment is generally achieved by adjusting the normal velocity increment. Let the z-direction safe distance between the mission satellite and the target satellite be Δz r , then the normal pulse amount applied in the first maneuver is:

[0147]

[0148] The mission star reaches D1 and applies a reverse pulse

[0149] Δv T3 =-(Δv T1 +Δv T2 ). (37)

[0150] The specific steps are as follows:

[0151] S021 determines the transfer time, n and n1, of the mission satellite approaching the target satellite;

[0152] S022 Calculate the second maneuvering track pulse Δv T2 ;

[0153] S023 Calculate the first maneuvering track pulse Δv T1 ;

[0154] S024 Calculate the normal pulse Δv of the first maneuver W ;

[0155] S025 Calculate the third maneuvering track pulse Δv T3 ;

[0156] S026 calculates the results according to the above steps, and applies the first maneuvering pulse (including the tracking pulse Δv) at time 0 D0. T1 and the normal pulse Δv W ), in n1T d Apply the second maneuvering pulse Δv at time T2 , arrive near the target star D1 nT d Apply the third maneuvering pulse Δv at time T3 . Fly around D1 to turn on the onboard laser rangefinder and perform error correction.

[0157] Step 3: After the mission satellite approaches from a long distance, eccentricity error correction is performed at D1;

[0158] Due to the high-order perturbations of the Earth and the space environment, there is a large eccentricity deviation during the close approach process, and further eccentricity correction is required when flying around D1. x ,e y The influence is different from that of the latitude argument u of the current mission star. d It is related to, and will change the semi-major axis difference. Suppose the eccentricity error of the mission star after approaching from a long distance is δe x ,δe y Therefore, the D1 correction is performed in four steps, with the latitude arguments separated by 90°, to correct the eccentricity difference while offsetting the positive and negative changes in the semi-major axis difference.

[0159] The specific steps for eccentricity error correction are:

[0160] S031, measurement mission satellite in u d = 0, the components of the eccentricity error on the x and y axes are δe x ,δe y ;

[0161] S032, calculate the components of the total pulse on the x and y axes required for eccentricity error correction, which are Δv Tex , ΔvTey :

[0162]

[0163]

[0164] S033, in u d = 0°, apply correction pulse If this is the first correction, combine this correction pulse with the third pulse mentioned in step 2 otherwise,

[0165] S034, in u d =90°, apply correction pulse

[0166] S035, in u d =180°, apply correction pulse

[0167] S036, in u d =270°, apply correction pulse

[0168] S037, measurement mission star u d = 0° eccentricity error, if it is not zero, execute step S032, otherwise, execute step S038;

[0169] S038, after completing the correction, the mission star is in u d = 0° when applying correction pulse Δv at3 To offset the atmospheric perturbations, the aircraft continued to fly around D1 for one cycle, and then made preparations for a close approach;

[0170] Δv at3 The calculation formula is

[0171]

[0172] Where n e To correct the total period of the flyby at D1.

[0173] Step 4: The mission satellite further adjusts the semi-major axis and eccentricity at D1 to prepare for hovering;

[0174] Set the hovering position to be L behind the target star B , left side L L , below L D First, we need to eliminate the anti-collision safety configuration in both the x and z directions at the same time. d =0°,u d=180°The fourth and fifth orbit control pulse maneuvers are carried out at two moments. The fourth maneuver has a pulse quantity Δv T4 and the fifth maneuver trace pulse Δv T5 The calculation method is as follows:

[0175]

[0176]

[0177] Where, ΔL T5 The track displacement generated by the fifth maneuver

[0178]

[0179] ΔL T4 The track displacement caused by the advance offset of the semi-major axis in the fourth maneuver, in is the actual distance between the mission star and the target star during the fourth maneuver. The normal pulse of the fourth maneuver is -Δv W After adjusting the eccentricity, the mission star is d =0° applied pulse-(Δv T4 +Δv T5 ), so that the mission star's orbital altitude remains unchanged, and this pulse is merged into the mission star's sixth maneuvering track pulse Δv T6 .

[0180] Step 5: The mission satellite approaches the designated hovering position from D1 through a three-pulse maneuver.

[0181] According to the law of relative motion, the mission star always has a drift in the velocity direction relative to the target star, and cannot be stably maintained at any relative position, but this property can be used to make the mission star drift to the specified position. Therefore, it is necessary to make a phase adjustment at D1 to compensate for the drift distance in advance. When the x direction meets the distance requirement, the mission star trail drifts to the specified position L B According to u d =180°, the relative distance when hovering below is L D , and obtain the required tracking pulse:

[0182]

[0183] This pulse will produce a tracking drift of ΔL D for

[0184]

[0185] Considering the atmospheric drag, the semi-major axis is biased in advance, and the resulting track displacement ΔL is obtained. ats, then the phase adjustment needs to be compensated and the distance configured is

[0186] ΔL s =ΔL D +L B +ΔL ats , (44)

[0187] Therefore, the pulse size required for phase adjustment is calculated as

[0188]

[0189] Where η is the mission star angular velocity.

[0190] The mission star needs to set the distance L in the z direction L , add normal pulse to achieve left bias. The normal pulse calculation formula is:

[0191]

[0192] The mission star needs to be in u d = 0°, a normal pulse is applied. d =270°, the mission satellite can reach the designated position on the left. Then the sixth maneuver of the mission satellite will have a pulse value of Δv T6 , the seventh maneuver trace pulse Δv T7 and the normal pulse Δv W7 , the eighth maneuver trace pulse Δv T8 The calculation formula is:

[0193] Δv T6 =Δv p -Δv T4 -Δv T5

[0194] Δv T7 =-Δv p +Δv D ,Δv W7 =Δv Wp

[0195] Δv T8 =Δv D (47)

[0196] The specific steps are as follows:

[0197] S051, calculate the required tracking pulse Δv during hovering D ;

[0198] S052, calculate the pulse size Δv required for phase adjustment p ;

[0199] S053, calculate the left bias normal pulse ΔvWp ;

[0200] S054, calculate the mission star's sixth pulse, seventh track and normal pulse, and eighth pulse size Δv T6 ,Δv T7 ,Δv W7 ,Δv T8 ;

[0201] S055, calculate the results according to the above steps, and then d = 0°, the sixth motor pulse Δv is applied T6 , after one cycle, the seventh tracking and normal pulse Δv is applied T7 ,Δv W7 , on the mission star u d =180°The eighth pulse Δv is applied T8 . Then when u d =270°, the mission star can reach the designated position on the left.

[0202] The control system according to the above-mentioned anti-collision hovering track control method for non-cooperative targets includes:

[0203] The first module is used to establish the relative motion equation based on orbital root difference and the relative motion control equation based on Gaussian perturbation;

[0204] The second module is used to control the mission satellite to form an anti-collision configuration through pulse maneuvers, so that the mission satellite can approach the target satellite from a long distance to a short distance.

[0205] The third module is used to control the mission satellite to perform eccentricity error correction after approaching from a long distance;

[0206] The fourth module is used to control the mission satellite to adjust the semi-major axis and eccentricity through pulse maneuvers;

[0207] The fifth module is used to control the mission satellite to approach the designated hovering position through pulse maneuvers.

[0208] Example 1

[0209] In the embodiment, the mission satellite and the target orbit type are sun synchronous orbits, the descending node local time is 9:00AM, the orbit altitude is 610km, the target satellite mass is 6.28kg, and the windward area is 3.14m 2 The mission satellite has a mass of 130 kg and a frontal area of ​​1.047 m 2 The drag coefficient is 2.2. The atmospheric density adopts the United States Standard Atmosphere Model (USSA76). At the initial time t0, the mission satellite is in the same orbit as the target satellite, 50 km behind the target satellite, and the latitude argument of the mission satellite is 0 at the beginning of the first maneuver.

[0210] The method of the present invention is used to achieve the mission satellite to quickly approach the target satellite from D0 = 50 km to the distance D1 = 4 km behind the target satellite, and then hover and observe at the specified position. The hovering position is L behind the target satellite. B =1000m, left side L L =500m, bottom L D =400m, the hovering time is 60s. Let the target satellite be D0 in n = 10 orbital periods (T d ) approaches the target star D1, and the second maneuver is performed n1 = 1.5 orbital periods after the first maneuver. Then the time when the mission star applies the pulse and the corresponding pulse size are calculated as:

[0211] t0: Track pulse Δv T =-0.376m / s, normal pulse Δv W =0.541m / s;

[0212] t0+1.5T d : Track pulse Δv T =0.034m / s;

[0213] t0+10T d : Track pulse Δv T =0.278m / s;

[0214] t0+10.25T d : Track pulse Δv T =-0.048m / s;

[0215] t0+10.5T d : Track pulse Δv T =-0.141m / s;

[0216] t0+10.75T d : Track pulse Δv T =0.039m / s;

[0217] t0+11T d : Track pulse Δv T =-0.065m / s;

[0218] t0+12T d : Track pulse Δv T =0.056m / s, normal pulse Δv W =-0.541m / s;

[0219] t0+12.5T d : Track pulse Δv T =-0.144m / s;

[0220] t0+13T d : Track pulse Δv T =0.134m / s;

[0221] t0+14T d : Track pulse Δv T =-0.200m / s, normal pulse Δv W =0.541m / s;

[0222] t0+14.5T d : Track pulse Δv T =-0.122m / s.

[0223] Mission star at 14.75T d The mission satellite quickly approaches the relative trajectory from 50km to 4km. Figure 2 , the anti-collision safety configuration during rapid approach (xz plane) is shown in Figure 3 , the mission satellite 4km correction and relative orbit trajectory are shown in Figure 4 The mission satellite approaches the designated position at 4 km and hovers to observe the relative trajectory. Figure 5 .

[0224] Parts of the present invention that are not described in detail belong to the well-known technologies of those skilled in the art.

Claims

1. A collision-avoiding hovering trajectory control method for non-cooperative targets, characterized in that: include: Establish the relative motion equation based on orbital root difference and the relative motion control equation based on Gaussian perturbation; Control the mission satellite to form an anti-collision configuration through pulse maneuvers, and achieve approaching the target satellite from a long distance to a short distance; Control the mission satellite to make eccentricity error correction after approaching from a long distance; Control the mission satellite to adjust the semi-major axis and eccentricity through pulse maneuvers; Control the mission satellite to approach the designated hovering position through pulse maneuvers; The relative motion equation based on the orbital root difference is: Among them, the position coordinates of the mission satellite in the orbital reference coordinate system are (x, y, z); the orbital elements of the mission satellite are The orbital elements of the target star are The difference in orbital elements between the mission star and the target star, that is, the relative orbital elements, is: a d is the semi-major axis of the mission star, e d is the eccentricity of the mission star, i d is the orbital inclination of the mission star, Ω d is the right ascension of the mission star’s ascending node, ω d is the mission star's perigee argument, u d is the latitude argument of the mission satellite; a c is the semi-major axis of the target star, e c is the eccentricity of the target star, i c is the orbital inclination of the target star, Ω c is the right ascension of the ascending node of the target star, ω c is the target star's perigee argument, u c is the latitude argument of the target star; Δe x , Δe y are the relative eccentricity vectors Components in the x and y axis directions; Δi x ,Δi y Relative inclination vector Components in the x and y axis directions; u0 is the latitude argument of the mission satellite’s initial position; The relative eccentricity vector and the relative inclination vector are respectively: The relative motion control equation based on Gaussian perturbation is: Among them, for the mission satellite orbit control in a near-circular orbit, the change in the number of mission satellite orbit elements is [δa,δi,δΩ,δe x ,δe y ]; V is the orbital velocity of the mission star, ΔV R ,ΔV T ,ΔV W are the velocity increments in radial, track and normal directions respectively; The control mission satellite forms an anti-collision configuration through pulse maneuvers to approach the target satellite from a long distance to a short distance, including: The mission star and the target star are orbiting in the same orbit, with the semi-major axis a d =a c = a, the initial position of the mission star is at position D0 behind the target star, and the mission star approaches from position D0 to the distance D1 behind the target star, including: S021, determine the transfer time of the mission satellite to the target satellite: let the target satellite be transferred from D0 to the target satellite in n full orbit periods T d Approach the target star D1, and the second maneuver is performed n1 orbital periods after the first maneuver, where n1 is an odd multiple of 0.5; S022, calculate the second maneuver trajectory pulse Δv considering the atmospheric drag perturbation T2 for: Where Δx r It is the safe distance between the mission star and the target star in the x direction when the latitude argument of the target star is 0; Considering the atmospheric drag perturbation, the semi-major axis attenuation is estimated to obtain the pulse Δv required to offset the atmospheric drag. at2 for: ρ is the atmospheric density, C D is the drag coefficient of the spacecraft, ε d is the surface mass ratio of the mission star, ε c is the surface mass ratio of the target star; S023, calculate the first maneuvering track pulse Δv T1 : Where, ΔL T1 The pulse Δv is taken into account to offset the atmospheric resistance. at1 The resulting displacement, ΔL T2 is the track displacement generated by the second maneuver, and the calculation formula is as follows: ΔL T2 =-3(n-n1)T d Δv T2 ; S024, calculate the normal pulse Δv of the first maneuver W : Where Δz r is the safe distance between the mission star and the target star in the z direction; S025, calculate the third maneuver track pulse quantity Δv T3 :Δv T3 =-(Δv T1 +Δv T2 ); S026, the mission satellite applies the first maneuver pulse at time 0 D0, and the first maneuver pulse includes a tracking pulse Δv T1 and the normal pulse Δv W , in n1T d Apply the second maneuvering pulse Δv at time T2 , arrive at the vicinity of the target star D1 at nT d Apply the third maneuvering pulse Δv at time T3 ; Fly around D1, turn on the onboard laser rangefinder and perform error correction.

2. The method for controlling a non-cooperative target-oriented collision-avoiding hovering trajectory according to claim 1, characterized in that: The control mission satellite performs eccentricity error correction after approaching from a long distance, including: The mission satellite performs eccentricity correction when it flies around D1, including: S031, measurement mission satellite in u d = 0°, the components of the eccentricity error on the x and y axes are δe x ,δe y ; S032, calculate the components of the total pulse on the x and y axes required for eccentricity error correction, which are Δv Tex , Δv Tey : S033, in u d = 0°, apply correction pulse If it is the first correction, then combine this correction pulse with the third maneuvering track pulse. otherwise, S034, in u d =90°, apply correction pulse S035, in u d =180°, apply correction pulse S036, in u d =270°, apply correction pulse S037, measurement mission satellite in u d = 0° eccentricity error, if it is not zero, execute step S032, otherwise, execute step S038; S038, after completing the correction, the mission star is in u d = 0° when applying correction pulse Δv at3 To offset the atmospheric perturbations, the aircraft continued to fly around D1 for one cycle, and then made preparations for a close approach; Δv at3 The calculation formula is: Among them, n e To correct the total period of the flyby at D1.

3. The method for controlling a non-cooperative target-oriented collision-avoiding hovering trajectory according to claim 2, characterized in that: The control mission satellite adjusts the semi-major axis and eccentricity through pulse maneuvers, including: The fixed hovering position of the mission star is L behind the target star. B , left side L L , below L D Position, respectively at u d =0°, u d =180°The fourth and fifth orbital control pulse maneuvers were carried out at two moments; After adjusting the eccentricity, the mission star is d =0° applied pulse-(Δv T4 +Δv T5 ), so that the mission star's orbital altitude remains unchanged, and this pulse is merged into the mission star's sixth maneuvering track pulse Δv T6 .

4. The method for controlling a non-cooperative target-oriented collision-avoiding hovering trajectory according to claim 3, characterized in that: The fourth maneuver track pulse amount Δv of the fourth and fifth orbit control pulse maneuvers T4 and the fifth maneuver trace pulse Δv T5 The calculation formula is as follows: Where, ΔL T5 is the track displacement produced by the fifth maneuver, ΔL T4 The track displacement caused by the advance offset of the semi-major axis in the fourth maneuver, in It is the actual distance between the mission star and the target star during the fourth maneuver.

5. The method for controlling a non-cooperative target-oriented collision-avoiding hovering trajectory according to claim 4, characterized in that: The control mission satellite approaches the designated hovering position through pulse maneuvers, including: S051, calculate the required tracking pulse Δv during hovering D : S052, calculate the pulse size Δv required for phase adjustment p : Among them, the semi-major axis offset displacement ΔL caused by offsetting atmospheric drag is considered ats , the phase adjustment needs to be compensated and configured with a distance of ΔL s =ΔL D +L B +ΔL ats ; Δv D The resulting track drift ΔL D for: η is the mission star angular velocity; S053, calculate the left bias normal pulse Δv Wp : S054, calculate the mission satellite's sixth maneuver track pulse Δv T6 , the seventh maneuver trace pulse Δv T7 and the normal pulse Δv W7 , the eighth maneuver trace pulse Δv T8 The calculation formula is: Δv T6 =Δv p -Δv T4 -Δv T5 Δv T7 =-Δv p +Δv D ,Δv W7 =Δv Wp ; Δv T8 =Δv D S055, in turn on the mission star u d = 0°, the sixth motor tracking pulse Δv is applied T6 , after one cycle, the seventh maneuvering track and normal pulse Δv is applied T7 ,Δv W7 , on the mission star u d =180°, the eighth motor tracking pulse Δv is applied T8 , then when u d =270°, the mission star reaches the designated position on the left.

6. The control system of the anti-collision hovering trajectory control method for non-cooperative targets according to any one of claims 1 to 5, characterized in that: include: The first module is used to establish the relative motion equation based on orbital root difference and the relative motion control equation based on Gaussian perturbation; The second module is used to control the mission satellite to form an anti-collision configuration through pulse maneuvers, so that the mission satellite can approach the target satellite from a long distance to a short distance. The third module is used to control the mission satellite to perform eccentricity error correction after approaching from a long distance; The fourth module is used to control the mission satellite to adjust the semi-major axis and eccentricity through pulse maneuvers; The fifth module is used to control the mission satellite to approach the designated hovering position through pulse maneuvers.

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