Long-distance rendezvous orbit maneuver calculation method and system based on multi-circle latitude argument

Abstract

The present invention relates to a method for calculating a long-distance rendezvous orbit maneuver based on multiple laps of latitude arguments, comprising: constructing a long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude arguments; performing a perturbation orbit multi-pulse maneuver calculation based on multiple laps of latitude arguments for given maneuvering parameters to obtain the relative position and velocity of the long-distance rendezvous terminal; using the multi-lap maneuvering point latitude arguments and maneuvering pulses as design variables, and the terminal's relative position and velocity as constraints, employing a sequential quadratic programming algorithm to stably solve the long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude arguments, thereby obtaining the maneuvering pulse design variable values, the number of laps, and the maneuvering point latitude arguments for the given number of laps. Accordingly, the present invention proposes a long-distance rendezvous orbit maneuver calculation system. The present invention overcomes the convergence instability problem caused by the fixed number of laps in traditional long-distance rendezvous orbit maneuver calculation technology, and achieves stable calculation of long-distance rendezvous orbit maneuvers by performing a cross-lap search using multiple laps of latitude arguments.

Description

Technical Field

[0001] The present invention belongs to the field of aerospace technology and relates to a long-distance rendezvous orbit maneuver calculation method and system based on multi-circle latitude argument, which can be applied to the long-distance rendezvous orbit maneuver calculation for manned spaceflight and on-orbit service. Background Art

[0002] Rendezvous and docking involves the simultaneous rendezvous of two spacecraft at the same orbital position and at the same or similar speeds, connecting them structurally into a single entity. This involves two spacecraft: a target and a pursuer. The target typically does not perform orbital maneuvers, while the pursuer flies toward the target through a series of orbital maneuvers.

[0003] In traditional rendezvous orbit maneuver calculations, when the relative distance is relatively close, time and maneuver pulses are used as design variables; when the relative distance is relatively far, the fixed maneuvering point orbital position (such as the perigee and apogee of a certain circle) is used as the pulse design variable, or the fixed number of circles is used as the maneuvering point latitude argument and pulse design variables. This type of technology has been continuously used in close-range and long-range rendezvous for manned spacecraft, cargo spacecraft, and space laboratories. However, with the continued use of low-latitude launch sites (such as Wenchang) in engineering missions, the perigee and apogee of the tracker after entering orbit are near the equator, and the calculation of the number of circles is often measured based on the ascending node in the equatorial plane. At the same time, orbital perturbations cause the perigee and apogee to drift near the ascending node. In this case, the traditional maneuver calculation technology that fixes the number of circles and only uses the maneuvering point latitude argument as the design variable has the problem of unstable convergence due to the change in the circle count. Summary of the Invention

[0004] The technical problem solved by the present invention is: to overcome the shortcomings of the existing technology and propose a long-distance rendezvous orbit maneuver calculation method and system based on multi-circle latitude argument, so as to overcome the problem of unstable convergence of traditional rendezvous orbit maneuver calculation technology caused by the maneuvering point position near the equator, and provide more stable orbit maneuver calculation support for long-distance rendezvous.

[0005] The solution of the present invention is:

[0006] A method for calculating long-distance rendezvous orbit maneuvers based on multi-circle latitude arguments, comprising:

[0007] Determine the long-range rendezvous orbit maneuver plan based on multiple laps of latitude arguments;

[0008] For given maneuvering parameters, a multi-pulse maneuver of the perturbation orbit is calculated based on the latitude arguments of multiple circles to obtain the relative position and velocity of the long-range rendezvous terminal.

[0009] Taking the multi-circle maneuvering point latitude argument and maneuvering pulse as design variables, and the relative position and velocity of the terminal as constraints, a sequential quadratic programming algorithm is used to stably solve the long-distance rendezvous orbit maneuver scheme based on multi-circle latitude argument. The maneuvering pulse design variable value, the number of circles, and the maneuvering point latitude argument under this number of circles are obtained.

[0010] Preferably, the long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments is as follows:

[0011] The first maneuver is performed along the track at the apogee of loop N1. t1 ;

[0012] The second maneuver was at the N2 latitude Execute pulse Δv along the normal direction z2 ;

[0013] The third maneuver is performed along the track at the perigee of the N3 loop. t3 ;

[0014] The fourth maneuver is at the N circle and latitude angle near the N4 circle

[0015] φ4∈[φ 4L ,φ 4U ]=[2Nπ+u 4min ,2(N+1)π+u4 max ] executes pulse Δv along the track t4 ;

[0016] Among them, φ 4L represents the lower limit of the latitude argument of the fourth orbital maneuver, φ 4U represents the upper limit of the latitude argument of the fourth orbital maneuver, N represents the number of orbital maneuvers, u 4min Indicates the lower limit of the latitude argument of the initial cycle of the fourth orbital maneuver, u 4max Indicates the upper limit of the latitude argument of the fourth orbital maneuvering terminal pass.

[0017] Preferably, after performing four orbital maneuvers on the N1, N2, N3 and Nth turns, at the terminal time t f Requirements: Relative position of the terminal in the relative coordinate system Relative position to aim Equal, terminal relative speed Relative speed to aim equal.

[0018] Preferably, the relative coordinate system refers to a right-handed coordinate system (track direction) with the target center of mass as the origin, the direction from the center of the earth to the target center of mass as the x-axis (radial), the normal along the orbital plane as the z-axis (normal), and the y-axis in the orbital plane forming a right-handed coordinate system (track direction).

[0019] Preferably, the perturbation orbit multi-pulse maneuver calculation based on multiple laps of latitude argument is performed to obtain the relative position and velocity of the long-distance rendezvous terminal as follows:

[0020] 2.1) Assume that the initial tracking time is t0. According to the tracking initial position r0 and velocity v0 in the J2000 coordinate system, the time t1 from the tracking initial position r0 to the apogee of the N1th circle, the position r1 and velocity v1 of the tracker when it reaches the apogee of the N1th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t1 Convert the pulse Δv1 in the J2000 coordinate system, apply Δv1 to the tracker, and obtain the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system;

[0021] 2.2) Based on the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the latitude argument u of the tracker from the position after the first maneuver to the N2 circle z2 The time at t2, position r2, and velocity v2 are converted to the normal execution pulse Δv in the relative coordinate system through coordinate transformation. z2 Convert the pulse Δv2 in the J2000 coordinate system, apply Δv2 to the tracker, and obtain the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system;

[0022] 2.3) Based on the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system, the time t3, position r3 and velocity v3 of the tracker from the position after the second maneuver to the perigee of the N3th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t3 Convert the pulse Δv3 in the J2000 coordinate system, apply Δv3 to the tracker, and obtain the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system;

[0023] 2.4) Based on the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system, the time t4, position r4 and velocity v4 of the tracker from the position after the third maneuver to the latitude argument φ4 of the Nx circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t4 Convert the pulse Δv4 in the J2000 coordinate system, apply Δv4 to the tracker, and obtain the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system;

[0024] 2.5) Based on the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the time from the fourth maneuver to the terminal time t f The tracking track is used to obtain the tracker terminal position r in the J2000 coordinate system. f , speed v f ;

[0025] 2.6) According to the initial position of the target in the J2000 coordinate system Initial velocity The perturbation orbit integral is used to calculate the target from the initial time to the terminal time t f The trajectory of the target is obtained in the J2000 coordinate system. Terminal velocity According to r f 、v f and Calculate the relative position of the tracker and the target Relative speed between the tracker and the target

[0026] Preferably, the implementation method of using the sequential quadratic programming algorithm to stably solve the long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude argument is as follows:

[0027] 3.1) With x=(Δv t1 ,u z2 ,Δv z2 ,Δv t3 ,φ4,Δv t4 ) is the design variable, the relative position and speed of the terminal are used as constraints, and the total speed increment is used as the objective function to construct a planning model; the objective function is f=|Δv t1 |+|Δv z2 |+|Δv t3 |+|Δv t4 |; The constraints on the relative position and velocity of the terminals satisfy

[0028] 3.2) A sequential quadratic programming algorithm is used to perform a stable search solution, in which a multi-pulse maneuver of the perturbation orbit is calculated for each set of design variables given by the algorithm, and the relative position of the tracker and the target is calculated. Relative speed between the tracker and the target The value of the objective function f is fed back to the sequential quadratic programming algorithm. If the value of a set of design variables corresponds to If the relative position and velocity constraints of the terminal are satisfied and the value of the objective function f is minimized, then this set of design variables is the solution result, and the values ​​of the maneuver pulse design variables and the latitude arguments during the second and fourth maneuvers are obtained.

[0029] A long-distance rendezvous orbit maneuvering calculation system based on multi-circle latitude argument includes a long-distance rendezvous orbit maneuvering scheme determination module, a relative position and velocity calculation module, and a scheme solving module;

[0030] Long-distance rendezvous orbit maneuvering scheme determination module: used to determine the long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments;

[0031] Relative position and velocity calculation module: For given maneuvering parameters, multi-pulse maneuver calculation of perturbation orbit is performed based on multiple latitude arguments to obtain the relative position and velocity of the long-range rendezvous terminal;

[0032] Solution solving module: Taking the multi-circle maneuvering point latitude argument and maneuvering pulse as design variables, and the terminal relative position and speed as constraints, a sequential quadratic programming algorithm is used to stably solve the long-distance rendezvous orbit maneuvering scheme based on multi-circle latitude argument, and obtain the maneuvering pulse design variable value, the number of circles, and the maneuvering point latitude argument under this number of circles.

[0033] Preferably, the long-distance rendezvous orbit maneuvering scheme determination module is used to determine a long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments. The long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments is as follows:

[0034] The first maneuver is performed along the track at the apogee of loop N1. t1 ;

[0035] The second maneuver was at the N2 latitude Execute pulse Δv along the normal direction z2 ;

[0036] The third maneuver is performed along the track at the perigee of the N3 loop. t3 ;

[0037] The fourth maneuver is at the N circle and latitude angle near the N4 circle

[0038] φ4∈[φ 4L ,φ 4U ]=[2Nπ+u 4min ,2(N+1)π+u 4max ] executes pulse Δv along the track t4 ;

[0039] Among them, φ 4L represents the lower limit of the latitude argument of the fourth orbital maneuver, φ 4U represents the upper limit of the latitude argument of the fourth orbital maneuver, N represents the number of orbital maneuvers, u 4min Indicates the lower limit of the latitude argument of the initial cycle of the fourth orbital maneuver, u 4maxIndicates the upper limit of the latitude argument of the fourth orbital maneuver terminal loop;

[0040] After executing four orbital maneuvers in the N1, N2, N3 and N circles, at the terminal time t f Requirements: Relative position of the terminal in the relative coordinate system Relative position to aim Equal, terminal relative speed Relative speed to aim equal.

[0041] Preferably, the relative position and speed calculation module is implemented as follows:

[0042] 2.1) Assume that the initial tracking time is t0. According to the tracking initial position r0 and velocity v0 in the J2000 coordinate system, the time t1 from the tracking initial position r0 to the apogee of the N1th circle, the position r1 and velocity v1 of the tracker when it reaches the apogee of the N1th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t1 Convert the pulse Δv1 in the J2000 coordinate system, apply Δv1 to the tracker, and obtain the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system;

[0043] 2.2) Based on the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the latitude argument u of the tracker from the position after the first maneuver to the N2 circle z2 The time at t2, position r2, and velocity v2 are converted to the normal execution pulse Δv in the relative coordinate system through coordinate transformation. z2 Convert the pulse Δv2 in the J2000 coordinate system, apply Δv2 to the tracker, and obtain the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system;

[0044] 2.3) Based on the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system, the time t3, position r3 and velocity v3 of the tracker from the position after the second maneuver to the perigee of the N3th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t3 Convert the pulse Δv3 in the J2000 coordinate system, apply Δv3 to the tracker, and obtain the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system;

[0045] 2.4) Based on the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system, the time t4, position r4 and velocity v4 of the tracker from the position after the third maneuver to the latitude argument φ4 of the Nx circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t4 Convert the pulse Δv4 in the J2000 coordinate system, apply Δv4 to the tracker, and obtain the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system;

[0046] 2.5) Based on the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the time from the fourth maneuver to the terminal time t f The tracking track is used to obtain the tracker terminal position r in the J2000 coordinate system. f , speed v f ;

[0047] 2.6) According to the initial position of the target in the J2000 coordinate system Initial velocity The perturbation orbit integral is used to calculate the target from the initial time to the terminal time t f The trajectory of the target is obtained in the J2000 coordinate system. Terminal velocity According to r f 、v f and Calculate the relative position of the tracker and the target Relative speed between the tracker and the target

[0048] Preferably, the solution solving module is implemented as follows:

[0049] 3.1) With x=(Δv t1 ,u z2 ,Δv z2 ,Δv t3 ,φ4,Δv t4 ) is the design variable, the relative position and speed of the terminal are used as constraints, and the total speed increment is used as the objective function to construct a planning model; the objective function is f=|Δv t1 |+|Δv z2 |+|Δv t3 |+|Δv t4 |; The constraints on the relative position and velocity of the terminals satisfy

[0050] 3.2) A sequential quadratic programming algorithm is used to perform a stable search solution, in which a multi-pulse maneuver of the perturbation orbit is calculated for each set of design variables given by the algorithm, and the relative position of the tracker and the target is calculated. Relative speed between the tracker and the target The value of the objective function f is fed back to the sequential quadratic programming algorithm. If the value of a set of design variables corresponds to If the relative position and velocity constraints of the terminal are satisfied and the value of the objective function f is minimized, then this set of design variables is the solution result, and the values ​​of the maneuver pulse design variables and the latitude arguments during the second and fourth maneuvers are obtained.

[0051] The beneficial effects of the present invention compared with the prior art are:

[0052] (1) The present invention uses multiple circles of latitude and angle to measure orbital position, which is not affected by the sudden change caused by integer changes such as the increase in the number of circles. The position measurement continuity is better, which is conducive to the search of gradient-based optimization algorithms such as sequential quadratic programming algorithms.

[0053] (2) The long-distance rendezvous orbit maneuver calculation method provided by the present invention can stably calculate the long-distance rendezvous maneuver parameters when the initial perigee and apogee of the tracker are near the equator. Through cross-circle search, the problem of unstable convergence of traditional rendezvous orbit maneuver calculation technology caused by the maneuvering point position near the equator can be overcome, providing more stable orbit maneuver calculation support for long-distance rendezvous. BRIEF DESCRIPTION OF THE DRAWINGS

[0054] Figure 1 This is a schematic diagram of the maneuvering position search interval based on multiple circles of latitude arguments;

[0055] Figure 2 This is a flow chart of the rendezvous orbit maneuver calculation based on multiple latitude arguments.

[0056] Figure 3 is the semi-major axis variation history of the long-range rendezvous tracker;

[0057] Figure 4 is the phase angle variation history of long-distance rendezvous;

[0058] Figure 5 It is the history of the change of the relative distance of the long-distance encounter;

[0059] Figure 6 It is the history of relative velocity changes during long-distance encounters. DETAILED DESCRIPTION

[0060] The present invention will be further described below with reference to the accompanying drawings.

[0061] The present invention proposes a rendezvous orbit maneuver calculation method using multi-circle latitude arguments as position design variables. Through cross-circle search, the problem of unstable convergence of traditional rendezvous orbit maneuver calculation technology caused by the maneuvering point position near the equator can be overcome, providing more stable orbit maneuver calculation support for long-distance rendezvous.

[0062] The present invention uses the multi-circle latitude argument of the maneuvering point as the position design variable, combines it with the pulse design variable, and comprehensively utilizes the perturbation orbit integration and numerical optimization technology to stably calculate the long-distance rendezvous orbit maneuvering parameters.

[0063] like Figure 2 As shown, the specific steps of the present invention are as follows:

[0064] Step 1: Determine the long-distance rendezvous orbit maneuvering plan based on multiple latitude arguments;

[0065] The long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude argument is as follows:

[0066] The first maneuver is performed along the track at the apogee of loop N1. t1 ;

[0067] The second maneuver was at the N2 latitude Execute pulse Δv along the normal direction z2 ;

[0068] The third maneuver is performed along the track at the perigee of the N3 loop. t3 ;

[0069] The fourth maneuver is in the N circle near the N4 circle, with a latitude argument φ4∈[φ 4L ,φ 4U ]=[2Nπ+u 4min ,2(N+1)π+u 4max ] executes pulse Δv along the track t4 ;

[0070] Among them, φ 4L represents the lower limit of the latitude argument of the fourth orbital maneuver, φ 4U represents the upper limit of the latitude argument of the fourth orbital maneuver, N represents the number of orbital maneuvers, u 4min Indicates the lower limit of the latitude argument of the initial cycle of the fourth orbital maneuver, u 4max Indicates the upper limit of the latitude argument of the fourth orbital maneuvering terminal pass. Figure 1 This is a schematic diagram of the maneuvering position search interval based on multiple circles of latitude arguments.

[0071] Step 2: For the given maneuvering parameters, perform a multi-pulse maneuver of the perturbation orbit based on the multi-circle latitude argument to obtain the relative position and velocity of the long-range rendezvous terminal;

[0072] The implementation is as follows:

[0073] 2.1) Assume that the initial tracking time is t0. According to the tracking initial position r0 and velocity v0 in the J2000 coordinate system, the time t1 from the tracking initial position r0 to the apogee of the N1th circle, the position r1 and velocity v1 of the tracker when it reaches the apogee of the N1th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t1 Convert the pulse Δv1 in the J2000 coordinate system, apply Δv1 to the tracker, and obtain the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system;

[0074] The relative coordinate system takes the center of mass of the target as its origin, the direction from the center of the earth to the center of mass of the target as the x-axis (radial), the normal direction along the orbital plane as the z-axis (normal), and the y-axis in the orbital plane forms a right-handed coordinate system (track direction).

[0075] 2.2) Based on the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the latitude argument u of the tracker from the position after the first maneuver to the N2 circle z2 The time at t2, position r2, and velocity v2 are converted to the normal execution pulse Δv in the relative coordinate system through coordinate transformation. z2 Convert the pulse Δv2 in the J2000 coordinate system, apply Δv2 to the tracker, and obtain the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system;

[0076] 2.3) Based on the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system, the time t3, position r3 and velocity v3 of the tracker from the position after the second maneuver to the perigee of the N3th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t3 Convert the pulse Δv3 in the J2000 coordinate system, apply Δv3 to the tracker, and obtain the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system;

[0077] 2.4) Based on the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system, the time t4, position r4 and velocity v4 of the tracker from the position after the third maneuver to the latitude argument φ4 of the Nx circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t4 Convert the pulse Δv4 in the J2000 coordinate system, apply Δv4 to the tracker, and obtain the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system;

[0078] 2.5) Based on the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the time from the fourth maneuver to the terminal time t f The tracking track is used to obtain the tracker terminal position r in the J2000 coordinate system. f , speed v f ;

[0079] 2.6) According to the initial position of the target in the J2000 coordinate system Initial velocity The perturbation orbit integral is used to calculate the target from the initial time to the terminal time t f The trajectory of the target is obtained in the J2000 coordinate system. Terminal velocity According to r f 、v f and Calculate the relative position of the tracker and the target Relative speed between the tracker and the target

[0080] Step 3: Using the multi-circle maneuvering point latitude argument and maneuvering pulse as design variables, and the terminal relative position and velocity as constraints, a sequential quadratic programming algorithm is used to stably solve the long-distance rendezvous orbit maneuver scheme based on multi-circle latitude argument, and the maneuvering pulse design variable value, the number of circles, and the maneuvering point latitude argument under this number of circles are obtained.

[0081] The implementation is as follows:

[0082] 3.1) With x=(Δv t1 ,u z2 ,Δv z2 ,Δv t3 ,φ4,Δv t4 ) is the design variable, the relative position and speed of the terminal are used as constraints, and the total speed increment is used as the objective function to construct a planning model; the objective function is f=|Δv t1 |+|Δv z2 |+|Δv t3 |+|Δv t4 |; The constraints on the relative position and velocity of the terminals are satisfied

[0083]

[0084]

[0085] in, To aim at relative position, The relative speed for aiming.

[0086] 3.2) Use the sequential quadratic programming algorithm to perform a stable search solution, in which the values ​​of each set of design variables given by the algorithm are calculated in step 2 to calculate the relative position of the tracker and the target. Relative speed between the tracker and the target The value of the objective function f is fed back to the sequential quadratic programming algorithm. If the value of a set of design variables corresponds to If the relative position and velocity constraints of the terminal are satisfied and the value of the objective function f is minimized, then this set of design variables is the solution result, and the values ​​of the maneuver pulse design variables and the latitude arguments during the second and fourth maneuvers are obtained.

[0087] On the basis of the above, the present invention proposes a long-distance rendezvous orbit maneuvering calculation system based on multi-circle latitude argument, including a long-distance rendezvous orbit maneuvering scheme determination module, a relative position and speed calculation module and a scheme solving module.

[0088] Long-distance rendezvous orbit maneuvering scheme determination module: used to determine the long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments.

[0089] Relative position and velocity calculation module: For given maneuvering parameters, multi-pulse maneuver calculation of perturbation orbit is performed based on multi-circle latitude arguments to obtain the relative position and velocity of the long-distance rendezvous terminal.

[0090] Solution solving module: Taking the multi-circle maneuvering point latitude argument and maneuvering pulse as design variables, and the terminal relative position and speed as constraints, a sequential quadratic programming algorithm is used to stably solve the long-distance rendezvous orbit maneuvering scheme based on multi-circle latitude argument, and obtain the maneuvering pulse design variable value, the number of circles, and the maneuvering point latitude argument under this number of circles.

[0091] Among them, the long-distance rendezvous orbit maneuvering scheme determination module is used to determine the long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments. The long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments is as follows:

[0092] The first maneuver is performed along the track at the apogee of loop N1. t1 ;

[0093] The second maneuver was at the N2 latitude Execute pulse Δv along the normal direction z2 ;

[0094] The third maneuver is performed along the track at the perigee of the N3 loop. t3 ;

[0095] The fourth maneuver is at the N circle and latitude angle near the N4 circle

[0096] φ4∈[φ 4L,φ 4U ]=[2Nπ+u4 min ,2(N+1)π+u4 max ] executes pulse Δv along the track t4 ;

[0097] Among them, φ 4L represents the lower limit of the latitude argument of the fourth orbital maneuver, φ 4U represents the upper limit of the latitude argument of the fourth orbital maneuver, N represents the number of orbital maneuvers, u 4min Indicates the lower limit of the latitude argument of the initial cycle of the fourth orbital maneuver, u 4max Indicates the upper limit of the latitude argument of the fourth orbital maneuver terminal loop;

[0098] After executing four orbital maneuvers in the N1, N2, N3 and N circles, at the terminal time t f Requirements: Relative position of the terminal in the relative coordinate system Relative position to aim Equal, terminal relative speed Relative speed to aim equal.

[0099] The relative position and speed calculation module is implemented as follows:

[0100] 2.1) Assume that the initial tracking time is t0. According to the tracking initial position r0 and velocity v0 in the J2000 coordinate system, the time t1 from the tracking initial position r0 to the apogee of the N1th circle, the position r1 and velocity v1 of the tracker when it reaches the apogee of the N1th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t1 Convert the pulse Δv1 in the J2000 coordinate system, apply Δv1 to the tracker, and obtain the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system;

[0101] 2.2) Based on the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the latitude argument u of the tracker from the position after the first maneuver to the N2 circle z2 The time at t2, position r2, and velocity v2 are converted to the normal execution pulse Δv in the relative coordinate system through coordinate transformation. z2 Convert the pulse Δv2 in the J2000 coordinate system, apply Δv2 to the tracker, and obtain the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system;

[0102] 2.3) Based on the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system, the time t3, position r3 and velocity v3 of the tracker from the position after the second maneuver to the perigee of the N3th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t3 Convert the pulse Δv3 in the J2000 coordinate system, apply Δv3 to the tracker, and obtain the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system;

[0103] 2.4) Based on the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system, the time t4, position r4 and velocity v4 of the tracker from the position after the third maneuver to the latitude argument φ4 of the Nx circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t4 Convert the pulse Δv4 in the J2000 coordinate system, apply Δv4 to the tracker, and obtain the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system;

[0104] 2.5) Based on the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the time from the fourth maneuver to the terminal time t f The tracking track is used to obtain the tracker terminal position r in the J2000 coordinate system. f , speed v f ;

[0105] 2.6) According to the initial position of the target in the J2000 coordinate system Initial velocity The perturbation orbit integral is used to calculate the target from the initial time to the terminal time t f The trajectory of the target is obtained in the J2000 coordinate system. Terminal velocity According to r f 、v f and Calculate the relative position of the tracker and the target Relative speed between the tracker and the target

[0106] The solution solving module is implemented as follows:

[0107] 3.1) With x=(Δv t1 ,u z2 ,Δv z2 ,Δv t3 ,φ4,Δv t4) is the design variable, the relative position and speed of the terminal are used as constraints, and the total speed increment is used as the objective function to construct a planning model; the objective function is f=|Δv t1 |+|Δv z2 |+|Δv t3 |+|Δv t4 |; The constraints on the relative position and velocity of the terminals satisfy

[0108] 3.2) A sequential quadratic programming algorithm is used to perform a stable search solution, in which a multi-pulse maneuver of the perturbation orbit is calculated for each set of design variables given by the algorithm, and the relative position of the tracker and the target is calculated. Relative speed between the tracker and the target The value of the objective function f is fed back to the sequential quadratic programming algorithm. If the value of a set of design variables corresponds to If the relative position and velocity constraints of the terminal are satisfied and the value of the objective function f is minimized, then this set of design variables is the solution result, and the values ​​of the maneuver pulse design variables and the latitude arguments during the second and fourth maneuvers are obtained.

[0109] Example:

[0110] 1. Construct a rendezvous orbit maneuver plan based on multiple circles of latitude arguments.

[0111] 2. For given maneuvering parameters, accurate calculation of perturbation orbit multi-pulse maneuvers is achieved based on multi-circle latitude arguments to obtain the relative position and velocity of the long-range rendezvous terminal.

[0112] 3. Using the maneuvering point's multiple-turn latitude argument as the position design variable, combined with the maneuvering pulse design variable, and the terminal's relative position velocity as an equality constraint, a sequential quadratic programming algorithm is employed for a stable solution. The maneuvering pulse design variable value, the number of turns, and the maneuvering point's latitude argument for that number of turns are obtained.

[0113] The rendezvous orbit maneuver scheme based on multiple latitude arguments is constructed as follows:

[0114] The first maneuver is performed along the track at the apogee of N1 = 4 turns. t1 ;

[0115] The second maneuver is at N2 = 7 degrees of latitude Execute a pulse Δv along the normal direction z2 ;

[0116] The third maneuver is performed along the track at the perigee of N3 = 10 t3 ;

[0117] The fourth maneuver is performed along the track pulse Δv at the latitude argument φ4∈[26π+6,28π+0.35] near the apogee of the initial circle N4=13 t4 ;

[0118] At the terminal time t f = relative position of 86400s speed Relative position to aim speed Equal, where the relative position of the aiming is The relative speed of aiming is

[0119] Step 2 is implemented as follows:

[0120] 2.1) The initial time t0 = 0 corresponds to 2021-01-01 10:00:00.00 Beijing time. The initial tracking position r0 of the J2000 system is (-1553236.2509m, -6340792.4586m, 386023.4677m), the initial velocity v0 is (5615.371764m / s, -1694.784111m / s, -5246.116308m / s), and the corresponding classical orbital elements (semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, periapsis, true anomaly) are (6646000m, 0.016, 42.01°, 80.0°, 175.0°, 359.94°), using the perturbation orbit integral, the time from the initial position velocity to the fourth circle apogee is calculated as t1 = 18846.0569s, the position is (1599322.0713m, 6540591.8600m, -246480.3519m), and the velocity is (-5472.329421m / s, 1530.436670m / s, 5103.629084m / s), and the tracking pulse Δv is executed. t1 =27.444444 m / s The J2000 pulse Δv1 = (-19.663446 m / s, 5.499241 m / s 18.338614 m / s) is calculated by coordinate transformation (well known in the aerospace field). Applying Δv1 yields the position r1 and velocity v1 + Δv1 of the tracker after the first maneuver.

[0121] 2.2) Based on the position r1 and velocity v1+Δv1 of the tracker after the first maneuver, the perturbation orbit integral is used to calculate the latitude argument u of the tracker from the position after the first maneuver to the 7th circle 2z=258.4833° is t2=33648.2220s, position r1 is (4460675.4719m,-2351298.2256m,-4373721.2207m), speed is (2796.078022m / s,7129.552733m / s,-1070.887110m / s), the normal pulse Δv z2 =-21.990740 m / s The J2000 pulse Δv2 = (-14.358379 m / s, 3.175137 m / s, -16.350782 m / s) is calculated by coordinate transformation. Applying Δv2, we obtain the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system.

[0122] 2.3) Based on the position r2 and velocity v2+Δv2 of the tracker after the second maneuver, the time from the position after the second maneuver to the perigee of the 10th orbit is calculated by using the perturbation orbit integral: t3 = 48762.4897s, the position r3 is (-1627368.0025m, -6430235.2703m, 90010.3267m), and the velocity v3 is (5588.835352m / s, -1487.391264m / s, -5212.557658m / s). t3 =3.771399 m / s The J2000 pulse Δv3 = (2.707213 m / s, -0.7204487 m / s, -2.524945 m / s) is calculated by coordinate transformation. Applying Δv3, we obtain the position r3 and velocity v3 + Δv3 of the tracker after the third maneuver in the J2000 coordinate system.

[0123] 2.4) Based on the position r3 and velocity v3+Δv3 of the tracker after the third maneuver, the time from the position after the third maneuver to the multi-circle latitude argument φ4=87.93353rad (13 circles 358.22°) is calculated by using the perturbation orbit integral: t4=67797.9111s, the position r4 is (1865186.6916m, 6485993.5110m, -140399.4422m), the velocity v4 is (-5434.748957m / s, 1676.921502m / s, 5123.191278m / s), and the tracking pulse Δv is calculated. t4 =34.945091m / s The J2000 pulse Δv4 = (-24.814166m / s, 7.642123m / s, 23.388341m / s) is calculated by coordinate transformation. Applying Δv4, we can obtain the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system.

[0124] 2.5) Based on the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver, the perturbation orbit integral is used to calculate the time from the fourth maneuver to the terminal time t f = 86400s tracking trajectory, and obtain the tracker terminal position r f is (-4826142.4051m,-3339494.1573m,3346374.5234m), speed v f is (2115.154327m / s,-6524.382871m / s,-3456.845337m / s);

[0125] 2.6) According to the initial position of the target speed The corresponding classical orbit elements (semi-major axis, eccentricity, orbit inclination, right ascension of ascending node, angular distance of periapsis, true anomaly) are (6764000m, 0.000, 42.00°, 80.00°, 0.00°, 270.00°). The perturbation orbit integral is used to calculate the target from the initial time to the terminal time t f The target trajectory is obtained to obtain the target terminal position Terminal velocity According to (r f ,v f )and Calculate the relative position of the tracker to the target: The relative speed is

[0126] Step 3 includes the following specific steps:

[0127] 3.1) With x=(Δv t1 ,u z2 ,Δv z2 ,Δv t3 ,φ4,Δv t4 ) is the design variable, and the relative position velocity obtained by integrating the perturbation orbit is equal to the aiming relative position velocity as the equality constraint: Take the total velocity increment as the objective function: f = |Δv t1 |+|Δv z2 |+|Δv t3 |+|Δv t4 |, build a planning model;

[0128] 3.2) Use the sequential quadratic programming algorithm to search and solve, where the value of each set of design variables given by the algorithm is perturbed using step 2 to integrate the orbit and calculate the relative position and velocity of the terminal. The value of the objective function f is fed back to the equality constraint of the sequential quadratic programming algorithm. If the relative position and velocity constraints of the terminal are satisfied and the value of the objective function f is minimized, then this set of design variables is the solution result, and the values ​​of the maneuver pulse design variables and the latitude arguments during the second and fourth maneuvers are obtained.

[0129] The solution is as follows: the value of the design variable x obtained by the sequential quadratic programming algorithm is (27.444444 m / s, 258.4833°, -21.990740 m / s, 3.771399 m / s, 358.2200°, 34.945091 m / s), and the value of the objective function f is 88.2 m / s. At the same time, the orbit elements of the long-range rendezvous tracker and the relative orbit change history data can be obtained. The orbit semi-major axis change history is shown in the attached figure. Figure 3 The phase angle (the difference between the target and tracker latitude arguments) change history is shown in the attached Figure 4 As shown in the attached figure, the relative distance change history is shown in the attached figure. Figure 5 The relative speed change history is shown in the attached Figure 6 shown.

[0130] The above results show that according to the initial six design variables, the sequential quadratic algorithm is used. After the optimization is completed, the specific parameters of the six design variables are obtained, the final result is obtained, and the terminal constraints are met with the minimum energy.

[0131] Unlike traditional rendezvous orbit maneuver calculation techniques that use a fixed number of turns, this method uses multiple turns of latitude and argument to perform cross-turn searches, achieving stable maneuver calculations. This overcomes the convergence instability problem associated with fixed turns in traditional long-distance rendezvous orbit calculation techniques, enabling stable long-distance rendezvous orbit calculations through cross-turn searches using multiple turns of latitude and argument.

[0132] The contents not described in detail in the specification of the present invention belong to the common knowledge of professionals in this field.

Claims

1. A method for calculating long-distance rendezvous orbit maneuvers based on multi-circle latitude arguments, characterized in that include: Determine the long-range rendezvous orbit maneuver plan based on multiple laps of latitude arguments; For given maneuvering parameters, a multi-pulse maneuver of the perturbation orbit is calculated based on the latitude arguments of multiple circles to obtain the relative position and velocity of the long-range rendezvous terminal. Using the latitude argument and pulse of the maneuvering point over multiple turns as design variables, and the relative position and velocity of the terminal as constraints, a sequential quadratic programming algorithm is used to stably solve the long-distance rendezvous orbit maneuver scheme based on the latitude argument of multiple turns. The maneuvering pulse design variable value, the number of turns, and the latitude argument of the maneuvering point under this number of turns are obtained. The long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude argument is as follows: The first maneuver is performed along the track at the apogee of loop N1. t1 ; The second maneuver was at the N2 latitude Execute pulse Δv along the normal direction z2 ; The third maneuver is performed along the track at the perigee of the N3 loop. t3 ; The fourth maneuver is in the N circle near the N4 circle, with a latitude argument φ4∈[φ 4L ,φ 4U ]=[2Nπ+u 4min ,2(N+1)π+u 4max ] executes pulse Δv along the track t4 ; Among them, φ 4L represents the lower limit of the latitude argument of the fourth orbital maneuver, φ 4U represents the upper limit of the latitude argument of the fourth orbital maneuver, N represents the number of orbital maneuvers, u 4min Indicates the lower limit of the latitude argument of the initial cycle of the fourth orbital maneuver, u 4max Indicates the upper limit of the latitude argument of the fourth orbital maneuvering terminal pass.

2. The method for calculating long-distance rendezvous orbit maneuvers based on multi-circle latitude arguments according to claim 1, characterized in that: After executing four orbital maneuvers in the N1, N2, N3 and N circles, at the terminal time t f Requirements: Relative position of the terminal in the relative coordinate system Relative position to aim Equal, terminal relative speed Relative speed to aim equal.

3. The method for calculating long-distance rendezvous orbit maneuvers based on multiple latitude arguments according to claim 2, characterized in that: The relative coordinate system refers to a right-handed coordinate system with the center of mass of the target as the origin, the direction from the center of the earth to the center of mass of the target as the x-axis, the normal along the orbital plane as the z-axis, and the y-axis within the orbital plane.

4. The method for calculating long-distance rendezvous orbit maneuvers based on multiple latitude arguments according to claim 2, characterized in that: The method for calculating the perturbation orbit multi-pulse maneuver based on the multi-circle latitude argument to obtain the relative position and velocity of the long-range rendezvous terminal is as follows: 2.1) Assume that the initial tracking time is t0. According to the tracking initial position r0 and velocity v0 in the J2000 coordinate system, the time t1 from the tracking initial position r0 to the apogee of the N1th circle, the position r1 and velocity v1 of the tracker when it reaches the apogee of the N1th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t1 Convert the pulse Δv1 in the J2000 coordinate system, apply Δv1 to the tracker, and obtain the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system; 2.2) Based on the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the latitude argument u of the tracker from the position after the first maneuver to the N2 circle z2 The time at t2, position r2, and velocity v2 are converted to the normal execution pulse Δv in the relative coordinate system through coordinate transformation. z2 Convert the pulse Δv2 in the J2000 coordinate system, apply Δv2 to the tracker, and obtain the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system; 2.3) Based on the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system, the time t3, position r3 and velocity v3 of the tracker from the position after the second maneuver to the perigee of the N3th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t3 Convert the pulse Δv3 in the J2000 coordinate system, apply Δv3 to the tracker, and obtain the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system; 2.4) Based on the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system, the time t4, position r4 and velocity v4 of the tracker from the position after the third maneuver to the latitude argument φ4 of the Nx circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t4 Convert the pulse Δv4 in the J2000 coordinate system, apply Δv4 to the tracker, and obtain the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system; 2.5) Based on the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the time from the fourth maneuver to the terminal time t f The tracking track is used to obtain the tracker terminal position r in the J2000 coordinate system. f , speed v f ; 2.6) According to the initial position of the target in the J2000 coordinate system Initial velocity The perturbation orbit integral is used to calculate the target from the initial time to the terminal time t f The trajectory of the target is obtained in the J2000 coordinate system. Terminal velocity According to r f 、v f and Calculate the relative position of the tracker and the target Relative speed between the tracker and the target 5. The method for calculating long-distance rendezvous orbit maneuvers based on multiple latitude arguments according to claim 4, characterized in that: The implementation method of using the sequential quadratic programming algorithm to stably solve the long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude angle is as follows: 3.1) With x=(Δv t1 ,u z2 ,Δv z2 ,Δv t3 ,φ4,Δv t4 ) is the design variable, the relative position and speed of the terminal are used as constraints, and the total speed increment is used as the objective function to construct a planning model; the objective function is f=|Δv t1 |+|Δv z2 |+|Δv t3 |+|Δv t4 |; The constraints on the relative position and velocity of the terminals satisfy 3.2) A sequential quadratic programming algorithm is used to perform a stable search solution, in which a multi-pulse maneuver of the perturbation orbit is calculated for each set of design variables given by the algorithm, and the relative position of the tracker and the target is calculated. Relative speed between the tracker and the target The value of the objective function f is fed back to the sequential quadratic programming algorithm. If the value of a set of design variables corresponds to If the relative position and velocity constraints of the terminal are satisfied and the value of the objective function f is minimized, then this set of design variables is the solution result, and the values ​​of the maneuver pulse design variables and the latitude arguments during the second and fourth maneuvers are obtained.

6. A long-distance rendezvous orbit maneuvering calculation system based on multi-circle latitude argument, characterized in that: It includes a long-distance rendezvous orbit maneuvering scheme determination module, a relative position and speed calculation module, and a scheme solving module; Long-distance rendezvous orbit maneuvering scheme determination module: used to determine the long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments; Relative position and velocity calculation module: For given maneuvering parameters, multi-pulse maneuver calculation of perturbation orbit is performed based on multiple latitude arguments to obtain the relative position and velocity of the long-range rendezvous terminal; Solution Module: This module uses the latitude argument and pulse of the maneuvering point over multiple turns as design variables, and the relative position and velocity of the terminal as constraints. It employs a sequential quadratic programming algorithm to stably solve the long-distance rendezvous orbit maneuver scheme based on the latitude argument of multiple turns. The module then obtains the maneuvering pulse design variable value, the number of turns, and the latitude argument of the maneuvering point under that number of turns. The long-distance rendezvous orbit maneuvering scheme determination module is used to determine the long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments. The long-distance rendezvous orbit maneuvering scheme based on multiple circles of latitude arguments is as follows: The first maneuver is performed along the track at the apogee of loop N1. t1 ; The second maneuver was at the N2 latitude Execute pulse Δv along the normal direction z2 ; The third maneuver is performed along the track at the perigee of the N3 loop. t3 ; The fourth maneuver is in the N circle near the N4 circle, with a latitude argument φ4∈[φ 4L ,φ 4U ]=[2Nπ+u 4min ,2(N+1)π+u 4max ] executes pulse Δv along the track t4 ; Among them, φ 4L represents the lower limit of the latitude argument of the fourth orbital maneuver, φ 4U represents the upper limit of the latitude argument of the fourth orbital maneuver, N represents the number of orbital maneuvers, u 4min Indicates the lower limit of the latitude argument of the initial cycle of the fourth orbital maneuver, u 4max Indicates the upper limit of the latitude argument of the fourth orbital maneuver terminal loop; After executing four orbital maneuvers in the N1, N2, N3 and N circles, at the terminal time t f Requirements: Relative position of the terminal in the relative coordinate system Relative position to aim Equal, terminal relative speed Relative speed to aim equal.

7. The long-distance rendezvous orbit maneuvering calculation system based on multi-circle latitude argument according to claim 6, characterized in that: The relative position and speed calculation module is implemented as follows: 2.1) Assume that the initial tracking time is t0. According to the tracking initial position r0 and velocity v0 in the J2000 coordinate system, the time t1 from the tracking initial position r0 to the apogee of the N1th circle, the position r1 and velocity v1 of the tracker when it reaches the apogee of the N1th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t1 Convert the pulse Δv1 in the J2000 coordinate system, apply Δv1 to the tracker, and obtain the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system; 2.2) Based on the position r1 and velocity v1+Δv1 of the tracker after the first maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the latitude argument u of the tracker from the position after the first maneuver to the N2 circle z2 The time at t2, position r2, and velocity v2 are converted to the normal execution pulse Δv in the relative coordinate system through coordinate transformation. z2 Convert the pulse Δv2 in the J2000 coordinate system, apply Δv2 to the tracker, and obtain the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system; 2.3) Based on the position r2 and velocity v2+Δv2 of the tracker after the second maneuver in the J2000 coordinate system, the time t3, position r3 and velocity v3 of the tracker from the position after the second maneuver to the perigee of the N3th circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t3 Convert the pulse Δv3 in the J2000 coordinate system, apply Δv3 to the tracker, and obtain the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system; 2.4) Based on the position r3 and velocity v3+Δv3 of the tracker after the third maneuver in the J2000 coordinate system, the time t4, position r4 and velocity v4 of the tracker from the position after the third maneuver to the latitude argument φ4 of the Nx circle are calculated by using the perturbation orbit integral. The tracking execution pulse Δv in the relative coordinate system is converted by coordinate transformation. t4 Convert the pulse Δv4 in the J2000 coordinate system, apply Δv4 to the tracker, and obtain the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system; 2.5) Based on the position r4 and velocity v4+Δv4 of the tracker after the fourth maneuver in the J2000 coordinate system, the perturbation orbit integral is used to calculate the time from the fourth maneuver to the terminal time t f The tracking track is used to obtain the tracker terminal position r in the J2000 coordinate system. f , speed v f ; 2.6) According to the initial position of the target in the J2000 coordinate system Initial velocity The perturbation orbit integral is used to calculate the target from the initial time to the terminal time t f The trajectory of the target is obtained in the J2000 coordinate system. Terminal velocity According to r f 、v f and Calculate the relative position of the tracker and the target Relative speed between the tracker and the target 8. The long-distance rendezvous orbit maneuvering calculation system based on multi-circle latitude argument according to claim 7, characterized in that: The solution solving module is implemented as follows: 3.1) With x=(Δv t1 ,u z2 ,Δv z2 ,Δv t3 ,φ4,Δv t4 ) is the design variable, the relative position and speed of the terminal are used as constraints, and the total speed increment is used as the objective function to construct a planning model; the objective function is f=|Δv t1 |+|Δv z2 |+|Δv t3 |+|Δv t4 |; The constraints on the relative position and velocity of the terminals satisfy 3.2) A sequential quadratic programming algorithm is used to perform a stable search solution, in which a multi-pulse maneuver of the perturbation orbit is calculated for each set of design variables given by the algorithm, and the relative position of the tracker and the target is calculated. Relative speed between the tracker and the target The value of the objective function f is fed back to the sequential quadratic programming algorithm. If the value of a set of design variables corresponds to If the relative position and velocity constraints of the terminal are satisfied and the value of the objective function f is minimized, then this set of design variables is the solution result, and the values ​​of the maneuver pulse design variables and the latitude arguments during the second and fourth maneuvers are obtained.

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