A long-distance guidance real-time deviation prediction method based on a pulse guidance model

Abstract

The application provides a long-distance guidance real-time deviation prediction method based on a pulse guidance model, comprising the following steps: 1) obtaining instantaneous orbital elements of a tracker and instantaneous orbital elements of a target at time t0; 2) obtaining an autonomous rendezvous and docking guidance sequence autonomously planned by the tracker and not yet executed; 3) setting t f To predict the time end point, if the equivalent start time of the i-th guidance pulse is in the range of [t0-t f ], the method proceeds to step 4); otherwise, the method proceeds to step 8); 4) calculating the instantaneous orbital elements before the i-th orbit control; 5) calculating the change of the instantaneous orbital elements of the tracker caused by the guidance control; 6) calculating the instantaneous orbital elements of the tracker after the guidance pulse control; 7) the guidance pulse number i+1; and judging i, if i<=n, the method proceeds to step 4); otherwise, the method proceeds to step 8); 8) calculating a free flight orbit set; 9) calculating a relative trajectory in the orbit system of the target; and 10) calculating the long-distance guidance terminal deviation.

Description

Technical Field

[0001] This invention relates to a real-time deviation prediction method for long-range guidance based on a pulse guidance model, belonging to the field of space rendezvous and docking technology. Background Technology

[0002] The rapid rendezvous and docking long-range guidance trajectory consists of several free-flight and orbit control segments. During this process, due to factors such as dynamic model deviations, navigation deviations, and control deviations, the trajectory may deviate from the design values. The magnitude of the terminal deviation is influenced by a combination of factors. If the terminal deviation exceeds the normal design range, ground control measures are required to correct the trajectory or terminate the current mission to ensure spacecraft safety. Especially when the long-range guidance terminal is outside the telemetry and control area, ground control needs to assess risks based on trajectory predictions and promptly implement measures before the terminal leaves the control area. Failure to respond promptly can not only cause significant economic losses but may also jeopardize the safety of the space station. Summary of the Invention

[0003] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a real-time deviation prediction method for long-range guidance based on a pulse guidance model. This method solves the problem of rapid real-time prediction of relative trajectory and terminal deviation for long-range guidance during rapid rendezvous and docking, and provides a basis for flight control decision-making.

[0004] The technical solution of this invention is: a method for real-time deviation prediction of long-range guidance based on a pulse guidance model, comprising:

[0005] (1) Obtain the instantaneous orbital elements M of the tracker at time t0. c (t0) and the instantaneous orbital elements M of the target T (t0);

[0006] (2) Based on the on-orbit telemetry parameters of the rendezvous and docking tracker, obtain the autonomous rendezvous and docking guidance sequence that the tracker has autonomously planned but has not yet executed;

[0007] (3) Let t f To predict the end time, determine if the equivalent start-up time KT of the i-th guidance pulse is... i In [t0~t f If the result is within the specified range, proceed to step (4); otherwise, proceed to step (8).

[0008] (4) Extrapolate the tracker track to the equivalent start time of the next pulse, and calculate the instantaneous root track M before the i-th track control. c (KT i );

[0009] (5) Based on the instantaneous root of the track before the i-th track control M c (KT i ) and KTi The velocity increment V of the time-guided pulse i The instantaneous change in the track elements ΔM caused by guidance and control was calculated. c (KT i );

[0010] (6) Based on the instantaneous root of the track before the i-th track control, M c (KT i ) and the instantaneous change in the track elements ΔM c (KT i The instantaneous orbital elements M of the tracker after guidance pulse control are calculated. c ′(KT i );

[0011] (7) The number of guidance pulses is i+1; and i is judged. If i≤n, it means that the guidance sequence in the prediction interval has not been completed. Then let M c (t0)=M c ′(KT i If the condition is not met, proceed to step (4); otherwise, proceed to step (8).

[0012] (8) Calculate the set of free flight orbits based on the extrapolated starting point of the orbit;

[0013] (9) Calculate the relative trajectory of the target spacecraft in the orbital system based on the obtained set of free flight trajectories;

[0014] (10) The deviation of the long-distance guidance terminal is calculated based on the relative trajectory.

[0015] The instantaneous orbital elements M of the tracker at time t0 are obtained. c (t0) and the instantaneous orbital elements M of the target T (t0), including: obtaining the instantaneous orbital elements M of the tracker at time t0 based on the on-orbit telemetry parameters of the rendezvous and docking tracker. c (t0)=[a c e c i c Ω c ω c u c and the instantaneous orbital elements M of the target T (t0)=[a T e T i T Ω T ω T u T ]; where the subscript c represents the tracker and the subscript T represents the target; a, e, i, Ω, ω, u represent the semi-major axis of the orbit, eccentricity, inclination, right axis of the focus of ascent, argument of perigee, and argument of the orbit, respectively.

[0016] The tracker autonomously planned and yet-to-be-executed autonomous rendezvous and docking guidance sequence [(KT1,V1); ...; (KT...] n V n )]; where KT i The equivalent start-up time of the i-th guidance pulse is indicated; i is the number of guidance pulses, i = 1…n; n represents the total number of guidance pulses executed within the current prediction period.

[0017] The track track is extrapolated to the equivalent start-up time of the next pulse, and the instantaneous root M of the track before the i-th track control is calculated. c (KT i ), including: M c (t0) Perform orbit extrapolation calculations to obtain the set of orbital parameters [M] during the free flight time before tracker orbit control. c (t0),…,M c (KT i Let M be the instantaneous root of the track before the i-th track control. c (KT i )=[a, e, i, Ω, ω, u].

[0018] The instantaneous orbital element extrapolation method f(·) based on quasi-average elements is used to calculate M. c (t0) Perform orbit extrapolation calculations, i.e., M c (t)=f(M c (t0),t), t is in [t0~KT] i Within the range.

[0019] The instantaneous root of the track before the i-th track control is M. c (KT i ) and KT i The velocity increment V of the time-guided pulse i The instantaneous change in the track elements ΔM caused by guidance and control was calculated. c (KT i ),include:

[0020] The pulse control equation for a small eccentricity track is adopted, based on M c (KT i ) and V i =[ΔV t ,ΔV n ,ΔV r The instantaneous change in track element number ΔM caused by guidance and control was calculated. c (KT i )=[Δa,Δi,ΔΩ,Δe x ,Δe y ];

[0021]

[0022] Among them, Δa, Δi, ΔΩ, Δe x Δe y These represent the semi-major axis correction, orbital inclination correction, ascending node right ascension correction, and eccentricity correction in the x and y directions, respectively; Δv t Δv n Δv r These represent the components of the pulsed orbit control velocity increment along the track direction, normal direction, and radial direction in the tracker's orbital coordinate system, respectively; n is the average angular velocity. μ is the Earth's gravitational constant; an eccentricity vector is introduced. As a controlled orbital element in a near-circular orbit; where e x Along the nodal line of the orbital plane, pointing towards the ascending node, e y Within the orbital plane, perpendicular and ahead of e x .

[0023] The M c ′(KT i )=[a c ′、e c ′、i c ′、Ω c ′、ω c ′、u c ′]:

[0024]

[0025] The calculation of the free-flight trajectory set based on the extrapolation starting point includes: using an instantaneous orbital element extrapolation calculation method based on quasi-average elements to perform tracker orbital extrapolation calculation, obtaining the tracker's orbital parameters from the current point to the predicted endpoint; if the tracker executed a track control sequence within the prediction interval, the extrapolation starting point is the orbit M after the last track control. c ′(KT n ), calculate to obtain [KT n ~t f The set of free flight orbits within [M] c ′(KT n ),…,M c (t f Otherwise, the starting point for orbit extrapolation is M. c (t0), calculate to obtain [t0~t f The set of free flight trajectory parameters within [M] c (t0),…,M c (t f )).

[0026] The step of calculating the relative trajectory based on the obtained set of free flight trajectories includes:

[0027] The instantaneous orbital element extrapolation function f(·) is used to calculate M. T (t0) Perform orbit extrapolation calculations to obtain the target's prediction time interval [t0~t f The set of all orbital parameters within [M] T (t0),…,M T (t f )];

[0028] The calculated tracker orbits are stitched together to obtain the set of all orbital parameters [M] within the tracker's predicted time interval. c (t0),…,M c (t f )];

[0029] By [M] T (t0),…,M T (t f )] and [M c (t0),…,M c (t f )] Calculate the prediction time interval [t0~t f The relative motion trajectory of the internal target spacecraft in its orbital system [R] RVD (t0),…,R RVD (t f )).

[0030] Based on the relative trajectory, the deviation of the long-range guidance terminal is calculated, including: Let the time of the long-range guidance terminal be t. zd When t zd <t f Then, calculate the long-range guidance terminal deviation δR. RVD (t zd ) = R RVD (t zd )-R RVD_bc Among them, R RVD (t zd ) and R RVD_bc These are the predicted calculated value and the design nominal value of the relative motion state of the long-distance guidance terminal, respectively.

[0031] The advantages of this invention compared to the prior art are:

[0032] (1) The quasi-average root orbit extrapolation calculation method and pulse orbit control equation calculation method adopted in this invention realize rapid real-time prediction of relative trajectories for long-distance guidance. Compared with the orbit numerical recursive prediction calculation method based on differential equations, this method has clear physical concepts, low computational load, and fast operation speed; compared with the CW equation extrapolation calculation method, this method has high accuracy in the long-distance segment and can meet the prediction requirements within a relative distance of 100km. After multiple on-orbit verifications, this method is very suitable for real-time online trajectory analysis of multiple orbital maneuvering missions in a short period of time, applicable to various working conditions, greatly improving the real-time performance of flight control judgment and decision-making, and has broad application prospects.

[0033] (2) Long-range guidance orbit control deviations mainly include: equivalent start-up time, average orbit control attitude, and rendezvous and docking navigation. In the mathematical model of the guidance equations, these are reflected in the latitude argument deviation at start-up time, the magnitude and direction deviation of orbit control velocity increment, and the orbital elements. Based on mission characteristics, this invention can select to estimate the main deviations affecting the start-up time and velocity increment based on on-orbit telemetry, and incorporate them into the real-time prediction model to gradually improve prediction accuracy. This achieves high-precision prediction from the long-range guidance terminal, meeting the requirements for rapid judgment and decision-making in flight control missions. Attached Figure Description

[0034] Figure 1 This is a flowchart of the method of the present invention;

[0035] Figure 2 A flowchart illustrating the deviation prediction process for long-distance guidance terminals;

[0036] Figure 3 The relative trajectory diagram of the fast-moving distant guide. Detailed Implementation

[0037] This invention relates to a method for real-time deviation prediction in long-range guidance based on a pulse guidance model, comprising:

[0038] (1) Based on the on-orbit telemetry parameters of the rendezvous and docking tracker, obtain the instantaneous orbital elements M of the tracker at time t0. c (t0)=[a c e c i c Ω c ω c u c and the instantaneous orbital elements M of the target T (t0)=[a T e T i T Ω T ω T u T]; where the subscript c represents the tracker, and the subscript T represents the target; a, e, i, Ω, ω, u represent the semi-major axis of the orbit, eccentricity, inclination, right axis of the focus of ascent, argument of perigee, and argument of the orbit, respectively;

[0039] (2) Based on the on-orbit telemetry parameters of the rendezvous and docking tracker, obtain the autonomous fast rendezvous guidance sequence [(KT1,V1); ...; (KT...] that the tracker has autonomously planned but not yet executed. n V n )]; where KT i V represents the equivalent start-up time of the i-th guidance pulse; i KT i The velocity increment of the guidance pulse at any given time; i is the number of guidance pulses, i = 1…n;

[0040] (3) Let t f To predict the end time, determine if the equivalent start-up time KT of the i-th guidance pulse is... i In [t0~t f If the result is within the specified range, proceed to step (4); otherwise, proceed to step (8).

[0041] (4) The tracker track is pushed out to the next pulse equivalent power-on time:

[0042] The instantaneous orbital element extrapolation method f(·) based on quasi-average elements is adopted, and M is calculated according to a set step size. c (t0) Perform orbit extrapolation calculations, i.e., M c (t)=f(M c (t0),)t,t is in [t0~KT] i Within the range; the set of orbital parameters [M] during the free flight time before tracker orbit control is calculated. c (t0),…,M c (KT i Let M be the instantaneous root of the track before the i-th track control. c (KT i = [a, e, i, Ω, ω, u]; where, the algorithm for f(·) is found in the references [1] Gong Yulian, He Yingzi, Li Maomao et al. Derailment braking control method using instantaneous orbital element prediction [J]. Journal of Astronautics, 2020, 41(12): 1533-1543. [2] Sun Baoxiang, Precise orbit determination technology using GPS receiver [J], Space Control Technology and Application, 2009.10, 35(5): 20-25.

[0043] (5) Calculation of pulse guidance equation:

[0044] The pulsed orbit control equations for small eccentricity orbits are described in "Zhang Renwei. Satellite Orbital Attitude Dynamics and Control [M]. Beijing: Beijing University of Aeronautics and Astronautics Press, 1998"; [M] c (KT i ) and V i =[ΔV t ,ΔV n ,ΔV r Calculate the instantaneous change in track element number ΔM caused by guidance and control. c (KT i )=[Δa,Δi,ΔΩ,Δe x ,Δe y ].

[0045]

[0046] Among them, Δa, Δi, ΔΩ, Δe x Δe y These represent the semi-major axis correction, orbital inclination correction, ascending node right ascension correction, and eccentricity correction in the x and y directions, respectively; Δv t Δv n Δv r These represent the components of the pulsed orbital control velocity increment in the tracker's orbital coordinate system along the track direction (velocity direction within the orbital plane), the normal direction (parallel to the orbital plane normal direction), and the radial direction (along the geocentric direction). n is the average angular velocity. μ is the Earth's gravitational constant; to avoid singularities, an eccentricity vector is introduced. As a controlled orbital element in a near-circular orbit; e x Along the nodal line of the orbital plane, pointing towards the ascending node, e y Within the orbital plane, perpendicular and ahead of e x .

[0047] (6) The orbital elements are updated to the control parameters:

[0048] According to M c (KT i ) and ΔM c (KT i ), calculate the instantaneous orbital elements M of the tracker after guidance pulse control. c ′(KT i )=[a c ′、e c ′、i c ′、Ω c ′、ω c ′、u c ′]:

[0049]

[0050] (7) i = i + 1; if i ≤ n, it means that the guidance sequence within the prediction interval has not been completed, then let M c (t0)=M c ′(KT i If the condition is met, proceed to step (4); otherwise, proceed to step (8) to continue.

[0051] (8) The tracker orbit is extrapolated to the end of the predicted interval:

[0052] An instantaneous orbital element extrapolation method f(·) based on quasi-average elements is adopted. The tracker's orbit is extrapolated according to a set step size to obtain the track parameters from the current position to the predicted endpoint. If a track control sequence was executed within the tracker's prediction interval, the extrapolation starting point is the orbit M after the last track control. c ′(KT n The result obtained by calculation is [KT] n ~t f The set of free flight orbits within [M] c ′(KT n ),…,M c (t f Otherwise, the starting point for orbit extrapolation is M. c (t0), calculate to obtain [t0~t f The set of free flight trajectory parameters within [M] c (t0),…,M c (t f )).

[0053] (9) Relative trajectory calculation:

[0054] a. Using the instantaneous orbital element extrapolation function f(·), for M T (t0) Perform orbit extrapolation calculations to obtain the target's prediction time interval [t0~t f The set of all orbital parameters within [M] T (t0),…,M T (t f )).

[0055] b. Concatenate the tracker orbits calculated in steps (4) and (8) to obtain the set of all orbital parameters [M] within the tracker prediction time interval. c (t0),…,M c (t f )).

[0056] c. By [M] T (t0),…,M T (t f )] and [M c (t0),…,M c (tf )] Calculate the prediction time interval [t0~t f The relative motion trajectory of the internal target spacecraft in its orbital system [R] RVD (t0),…,R RVD (t f )).

[0057] (10) Let the time of the long-distance guidance terminal be t. zd If t zd <t f Then calculate the long-range guidance terminal deviation δR RVD (t zd ) = R RVD (t zd )-R RVD_bc Among them, R RVD (t zd ) and R RVD_bc These are the predicted calculated value and the design nominal value of the relative motion state of the long-distance guidance terminal, respectively.

[0058] The working process and working principle of the present invention will be further explained below with reference to specific embodiments.

[0059] To eliminate the residual risk of uncertainties and deviations during the long-range guidance phase of rapid rendezvous and docking missions, and addressing the need for contingency plans based on trajectory prediction during missions, a lightweight real-time prediction algorithm based on a pulse guidance model is proposed. The corresponding software runs on the integrated processing computer of the ground flight control support network. Based on on-orbit telemetry, it corrects model parameters and predicts the relative trajectory and terminal deviation during long-range guidance for rapid rendezvous and docking in real time, ensuring the successful completion of the Tianzhou cargo spacecraft's 3-hour rapid rendezvous and docking mission.

[0060] This invention provides a method for real-time deviation prediction in long-range guidance based on a pulse guidance model, the steps of which are as follows:

[0061] The rapid rendezvous and docking mission tracker is inserted into orbit, and nominal operating condition simulations are performed based on the inserted orbit to obtain the nominal relative rendezvous trajectory. During the long-range guidance segment of the rendezvous and docking, real-time telemetry-driven prediction algorithms run periodically, and all calculation methods employ... Figure 1 The algorithm shown is a "long-range guidance deviation prediction algorithm". The parameters used are automatically switched according to telemetry at different mission stages, gradually improving prediction accuracy. Based on mission requirements and computational load, the algorithm's execution cycle is determined to be 3 seconds.

[0062] Taking a 3-hour rapid rendezvous and docking mission as an example, the calculation process for predicting deviations of long-range guidance terminals is as follows: Figure 2 As shown, the specific steps are as follows:

[0063] (1) Tracker orbit insertion: The initial orbits of the tracker and target are obtained through telemetry. The orbit after orbit insertion deviation correction is introduced into the long-distance guidance deviation prediction calculation. Under the real-time telemetry drive, the relative trajectory of the long-distance guidance segment is periodically predicted and calculated.

[0064] (2) Calculation of phase modulation and out-of-plane correction pulse: The guidance parameters of the next phase modulation and out-of-plane correction pulse planned online by the tracker are obtained through telemetry. Together with the telemetry orbits of the two targets, the long-distance guidance deviation prediction calculation is introduced, and the relative trajectory of the long-distance guidance segment is periodically predicted and calculated.

[0065] (3) Calculate the first pulse of the integrated correction: Obtain the two integrated correction guidance parameters of the tracker's on-orbit planning through telemetry, and together with the actual orbits of the two targets, introduce the long-distance guidance deviation prediction calculation, and periodically predict and calculate the relative trajectory and long-distance guidance terminal deviation; During the execution of the first pulse of the integrated correction, evaluate the orbit control execution deviation parameters through telemetry, including the equivalent start-up time and the average attitude angle.

[0066] (4) Transition to comprehensive correction subpulse calculation: The comprehensive correction subpulse guidance parameters of the tracker's on-orbit planning are obtained through telemetry. Together with the actual orbits of the two targets and the guidance deviation parameters (including equivalent start-up time and average attitude angle) evaluated by the comprehensive correction first pulse execution process, the long-range guidance deviation prediction calculation is introduced, and the relative trajectory and long-range guidance terminal deviation are periodically predicted and calculated.

[0067] (5) Relative navigation establishment: After the rendezvous reaches the range of the relative navigation sensor, the spacecraft begins to output the relative motion state calculated by the relative navigation sensor via telemetry. The ground initiates a prediction algorithm based on the CW equation and compares it with the terminal prediction algorithm proposed in this invention.

[0068] If the predicted accuracy of the long-range guidance terminal meets the requirements for normal transition to short-range autonomous control, the mission continues; if the predicted deviation of the long-range guidance terminal exceeds the threshold and the simulation shows that the rendezvous mission can continue by modifying the control parameters, the corresponding control parameters are modified before leaving the station and the mission continues; if the predicted relative trajectory is unsafe and there is a risk of collision, the current mission is exited according to the command sent from the ground and collision avoidance measures are taken.

[0069] The relative trajectory and long-distance terminal accuracy predicted by a 3-hour rapid transit distance of approximately 100km are as follows: Figure 3 As shown, the effectiveness of the method proposed in this patent is verified, with a prediction error of less than ±1km in the trajectory and ±0.2km in the radial direction.

[0070] In the rapid docking mission of the Tianzhou cargo spacecraft, the long-range guidance real-time deviation prediction method based on the pulse guidance model provided by this project can realize rapid and high-precision real-time prediction of deviations of the remote guidance terminal, meet the needs of flight control prediction and strategy evaluation, and ensure the complete success of the mission.

[0071] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for predicting real-time deviations in long-range guidance based on a pulse guidance model, characterized in that, include: (1) Obtain the instantaneous orbital elements M of the tracker at time t0. c (t0) and the instantaneous orbital elements M of the target T (t0); (2) Based on the on-orbit telemetry parameters of the rendezvous and docking tracker, obtain the autonomous rendezvous and docking guidance sequence that the tracker has autonomously planned but has not yet executed; (3) Let t f To predict the end time, determine if the equivalent start-up time KT of the i-th guidance pulse is... i In [t0~t f If the result is within the specified range, proceed to step (4); otherwise, proceed to step (8). (4) Extrapolate the tracker track to the equivalent start time of the next pulse, and calculate the instantaneous root track M before the i-th track control. c (KT i ); (5) Based on the instantaneous root of the track before the i-th track control M c (KT i ) and KT i The velocity increment V of the time-guided pulse i The instantaneous change in the track elements ΔM caused by guidance and control was calculated. c (KT i ); (6) Based on the instantaneous root of the track before the i-th track control, M c (KT i ) and the instantaneous change in the track elements ΔM c (KT i The instantaneous orbital elements M of the tracker after guidance pulse control are calculated. c ′(KT i ); (7) The number of guidance pulses is i+1; and i is judged. If i≤n, it means that the guidance sequence in the prediction interval has not been completed. Then let M c (t0)=M c ′(KT i If the condition is not met, proceed to step (4); otherwise, proceed to step (8). (8) Calculate the set of free flight orbits based on the extrapolated starting point of the orbit; (9) Calculate the relative trajectory of the target spacecraft in the orbital system based on the obtained set of free flight trajectories; (10) The deviation of the long-distance guidance terminal is calculated based on the relative trajectory.

2. The method for real-time deviation prediction of long-range guidance based on a pulse guidance model according to claim 1, characterized in that, The instantaneous orbital elements M of the tracker at time t0 are obtained. c (t0) and the instantaneous orbital elements M of the target T (t0), including: obtaining the instantaneous orbital elements M of the tracker at time t0 based on the on-orbit telemetry parameters of the rendezvous and docking tracker. c (t0)=[a c e c i c Ω c ω c u c and the instantaneous orbital elements M of the target T (t0)=[a T e T i T Ω T ω T u T ]; where the subscript c represents the tracker and the subscript T represents the target; a, e, i, Ω, ω, u represent the semi-major axis of the orbit, eccentricity, inclination, right axis of the focus of ascent, argument of perigee, and argument of the orbit, respectively.

3. The method for real-time deviation prediction of long-range guidance based on a pulse guidance model according to claim 2, characterized in that, The tracker autonomously planned and yet-to-be-executed autonomous rendezvous and docking guidance sequence [(KT1,V1); ...; (KT...] n V n )]; where KT i The equivalent start-up time of the i-th guidance pulse is indicated; i is the number of guidance pulses, i = 1…n; n represents the total number of guidance pulses executed within the current prediction period.

4. The method for real-time deviation prediction of long-range guidance based on a pulse guidance model according to claim 3, characterized in that, The track track is extrapolated to the equivalent start-up time of the next pulse, and the instantaneous root M of the track before the i-th track control is calculated. c (KT i ), including: M c (t0) Perform orbit extrapolation calculations to obtain the set of orbital parameters [M] during the free flight time before tracker orbit control. c (t0),…,M c (KT i Let M be the instantaneous root of the track before the i-th track control. c (KT i )=[a, e, i, Ω, ω, u].

5. The method for real-time deviation prediction of long-range guidance based on a pulse guidance model according to claim 4, characterized in that, The instantaneous orbital element extrapolation method f(·) based on quasi-average elements is used to calculate M. c (t0) Perform orbit extrapolation calculations, i.e., M c (t)=f(M c (t0),t), t is in [t0~KT] i Within the range.

6. The method for real-time deviation prediction of long-range guidance based on a pulse guidance model according to claim 4, characterized in that, The instantaneous root of the track before the i-th track control is M. c (KT i ) and KT i The velocity increment V of the time-guided pulse i The instantaneous change in the track elements ΔM caused by guidance and control was calculated. c (KT i ),include: The pulse control equation for a small eccentricity track is adopted, based on M c (KT i ) and V i =[ΔV t ,ΔV n ,ΔV r The instantaneous change in track element number ΔM caused by guidance and control was calculated. c (KT i )=[Δa,Δi,ΔΩ,Δe x ,Δe y ]; Among them, Δa, Δi, ΔΩ, Δe x , Δe y These represent the semi-major axis correction, orbital inclination correction, ascending node right ascension correction, and eccentricity correction in the x and y directions, respectively; Δv t Δv n Δv r These represent the components of the pulsed orbit control velocity increment along the track direction, normal direction, and radial direction in the tracker's orbital coordinate system, respectively; n is the average angular velocity. μ is the Earth's gravitational constant; an eccentricity vector is introduced. As a controlled orbital element in a near-circular orbit; where e x Along the nodal line of the orbital plane, pointing towards the ascending node, e y Within the orbital plane, perpendicular and ahead of e x .

7. The method for real-time deviation prediction of long-range guidance based on a pulse guidance model according to claim 6, characterized in that, the M c ′(KT i )=[a c ′, e c ′, i c ′, Ω c ′, ω c ′, u c ′]:

8. The method for real-time deviation prediction of long-range guidance based on a pulse guidance model according to claim 6, characterized in that, The calculation of the free-flight trajectory set based on the extrapolation starting point includes: using an instantaneous orbital element extrapolation calculation method based on quasi-average elements to perform tracker orbital extrapolation calculation, obtaining the tracker's orbital parameters from the current point to the predicted endpoint; if the tracker executed a track control sequence within the prediction interval, the extrapolation starting point is the orbit M after the last track control. c ′(KT n ), calculate to obtain [KT n ~t f The set of free flight orbits within [M] c ′(KT n ),…,M c (t f Otherwise, the starting point for orbit extrapolation is M. c (t0), calculate to obtain [t0~t f The set of free flight trajectory parameters within [M] c (t0),…,M c (t f )).

9. A method for real-time deviation prediction of long-range guidance based on a pulse guidance model according to claim 8, characterized in that, The step of calculating the relative trajectory based on the obtained set of free flight trajectories includes: The instantaneous orbital element extrapolation function f(·) is used to calculate M. T (t0) Perform orbit extrapolation calculations to obtain the target's prediction time interval [t0~t f The set of all orbital parameters within [M] T (t0),…,M T (t f )]; The calculated tracker orbits are stitched together to obtain the set of all orbital parameters [M] within the tracker's predicted time interval. c (t0),…,M c (t f )]; By [M] T (t0),…,M T (t f )] and [M c (t0),…,M c (t f )] Calculate the prediction time interval [t0~t f The relative motion trajectory of the internal target spacecraft in its orbital system [R] RVD (t0),…,R RVD (t f )).

10. A method for real-time deviation prediction of long-range guidance based on a pulse guidance model according to claim 9, characterized in that, Based on the relative trajectory, the deviation of the long-range guidance terminal is calculated, including: Let the time of the long-range guidance terminal be t. zd When t zd <t f At that time, calculate the long-range guidance terminal deviation δR RVD (t zd ) = R RVD (t zd )-R RVD_bc Among them, R RVD (t zd ) and R RVD_bc These are the predicted calculated value and the design nominal value of the relative motion state of the long-distance guidance terminal, respectively.

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