An Online Trajectory Planning Method for a Multi-Orbit Satellite to Deploy a Space Transfer Vehicle

Abstract

The present invention discloses an online trajectory planning method for a multi-orbit satellite-deployed space transfer vehicle, including: receiving a satellite deployment task uploaded from the ground; determining whether the current remaining fuel meets the requirements of the satellite deployment task; if it is determined that the current remaining fuel meets the requirements of the satellite deployment task, then judging the task type according to the current mean orbital elements and the target mean orbital elements; wherein, the task type includes: altitude adjustment, inclination adjustment, and combined adjustment; calculating the ignition duration and the latitude argument of the ignition section according to the determined different task types, and generating an ignition sequence; performing satellite deployment planning according to the generated ignition sequence; transmitting the satellite deployment planning result to the ground, and if the ground allows execution, then performing the task according to the satellite deployment planning result. The purpose of the present invention is to enable the space transfer vehicle to have the ability of online autonomous trajectory planning and give full play to its strong flexibility.

Description

Technical Field

[0001] The invention belongs to the technical field of flight vehicle trajectory planning, and particularly relates to an online trajectory planning method for a space transfer vehicle for multi-orbit satellite deployment. Background Technique

[0002] The space transfer vehicle has the characteristics of multiple starts, long-term on-orbit operation, autonomous flight, and strong orbital maneuverability. It has strong flexibility and versatility and can adapt to various different mission situations.

[0003] For the space transfer vehicle, it can only execute flight missions according to the pre-loaded targets at present. When encountering special situations such as large deviations in the basic orbit injection or the need to temporarily change the target orbit, it is difficult to ensure the successful execution of the mission only by the flight program loaded before takeoff. Moreover, the ability of the space transfer vehicle to participate in multi-satellite launches and long-term on-orbit operation makes it have a very broad application prospect. For example, it can carry backup satellites and quickly transfer to the target orbit to complete satellite deployment when needed. Since the target orbit is unknown before takeoff for such missions, it is impossible to load the target values before launch.

[0004] Based on the above requirements, it is necessary to propose an online trajectory planning method for a space transfer vehicle for multi-orbit satellite deployment, so that the space transfer vehicle has the ability of online autonomous trajectory planning and gives full play to its strong flexibility. Summary of the Invention

[0005] The technical problem to be solved by the invention: Overcoming the deficiencies of the prior art, providing an online trajectory planning method for a space transfer vehicle for multi-orbit satellite deployment, aiming to enable the space transfer vehicle to have the ability of online autonomous trajectory planning and give full play to its strong flexibility.

[0006] In order to solve the above technical problems, the invention discloses an online trajectory planning method for a space transfer vehicle for multi-orbit satellite deployment, including:

[0007] Receiving the satellite deployment task uploaded from the ground;

[0008] Judging whether the current remaining fuel meets the requirements of the satellite deployment task;

[0009] If it is determined that the current remaining fuel meets the requirements of the satellite deployment task, then judging the task type according to the current mean orbital elements and the target mean orbital elements; wherein, the task type includes: altitude adjustment, inclination adjustment, and combined adjustment;

[0010] Calculating the ignition duration and the latitude argument of the ignition section according to the determined different task types, and generating an ignition sequence;

[0011] Conducting satellite deployment planning according to the generated ignition sequence;

[0012] Download the satellite deployment planning result to the ground. If the ground allows execution, execute the task according to the satellite deployment planning result.

[0013] In the above online trajectory planning method for the multi-orbit satellite deployment space transfer vehicle, determining whether the current remaining fuel meets the satellite deployment task requirements includes:

[0014] Calculate the required velocity Δv for inclination adjustment i :

[0015]

[0016] where μ represents the gravitational constant, represents the mean semi-major axis of the target orbit, e obj represents the instantaneous eccentricity of the target orbit, and Δi represents the inclination deviation between the current orbit and the target orbit;

[0017] Calculate the required velocity Δv for semi-major axis adjustment a :

[0018]

[0019]

[0020]

[0021] where, represents the mean eccentricity of the target orbit, r p represents the perigee radius of the transfer orbit, e2 represents the mean eccentricity of the transfer orbit, represents the mean semi-major axis of the transfer orbit, a p represents the mean semi-major axis of the current orbit, e p represents the mean eccentricity of the current orbit;

[0022] Determine the total velocity increment Δv: Δv = [Δv i Δv a T ;

[0023] According to the total velocity increment Δv, calculate the required fuel m zy :

[0024]

[0025] where m0 represents the current vehicle mass, V e represents the jet velocity of the engine;

[0026] According to m zy , determine whether the current remaining fuel meets the satellite deployment task requirements; among them, if m zyIf it is less than the current remaining fuel, it is determined that the current remaining fuel meets the requirements of the satellite deployment mission; otherwise, it does not meet the requirements.

[0027] In the above online trajectory planning method for a multi-orbit satellite deployment space transfer vehicle, according to the current mean orbital elements and the target mean orbital elements, the mission type is judged, including:

[0028] If Then it is determined that the mission type is altitude adjustment;

[0029] If Then it is determined that the mission type is inclination adjustment;

[0030] If And Then it is determined that the mission type is combined adjustment;

[0031] Wherein, represents the target orbit mean orbital inclination, i p represents the current orbit mean orbital inclination, a limit represents the discrimination threshold for altitude adjustment, i limit represents the discrimination threshold for inclination adjustment.

[0032] In the above online trajectory planning method for a multi-orbit satellite deployment space transfer vehicle, according to the determined different mission types, the ignition duration and the ignition section latitude amplitude angle are calculated, and an ignition sequence is generated, including:

[0033] If the mission type is altitude adjustment, then there are:

[0034]

[0035]

[0036]

[0037]

[0038] Wherein, represents the engine fuel consumption per second, t1 represents the duration of the first ignition time, Δv A represents the required velocity increment for altitude adjustment at the first ignition, t2 represents the duration of the second ignition time, Δv C represents the required velocity increment for altitude adjustment at the second ignition, represents the latitude amplitude angle at the start time of the first ignition, u start represents the mean latitude amplitude angle at the execution time, represents the latitude amplitude angle at the start time of the second ignition, n0 represents the current orbit angular velocity, n t represents the transfer orbit angular velocity;

[0039] If the task type is inclination adjustment, then there is:

[0040]

[0041]

[0042] If the task type is combined adjustment, then there is:

[0043]

[0044]

[0045]

[0046]

[0047] In the above online trajectory planning method for deploying a space transfer vehicle by a multi-orbit satellite, it further includes: using the mean orbital elements [a p e p i p Ω p ω p u0] at the current moment t0 as the initial orbital elements, performing one orbit recurrence to obtain the mean orbital elements [a start at the task execution moment t start e start i start Ω start ω start u start ]; where, Ω p represents the mean right ascension of the ascending node of the current orbit, ω p represents the mean argument of perigee of the current orbit, and u0 represents the mean argument of latitude of the current orbit; a start represents the mean semi-major axis at the execution moment, e start represents the mean eccentricity at the execution moment, i start represents the mean orbital inclination at the execution moment, Ω start represents the mean right ascension of the ascending node at the execution moment, ω start represents the mean argument of perigee at the execution moment.

[0048] In the above online trajectory planning method for deploying a space transfer vehicle by a multi-orbit satellite, according to the generated ignition sequence, satellite deployment planning is performed, including:

[0049] According to the generated ignition sequence, calculate the starting moment of the tracking and control arc segment search and the starting argument of latitude of the tracking and control arc segment search, and perform the tracking and control arc segment search; where, the starting moment of the tracking and control arc segment search is the moment when the orbit transfer is completed;

[0050] Determine the latitude argument of the satellite deployment according to the ground station measurement and control conditions, so that the satellite deployment process is within the measurement and control range of the ground station.

[0051] In the above online trajectory planning method for the space transfer vehicle in the multi-orbit satellite deployment, the calculation processes of the starting time of the measurement and control arc segment search and the starting latitude argument of the measurement and control arc segment search are as follows:

[0052] If the mission type is altitude adjustment, then:

[0053]

[0054]

[0055]

[0056]

[0057] Among them, represents the eccentric anomaly at the end of the first ignition, e tr represents the eccentricity of the transfer orbit, represents the latitude argument at the end of the first ignition, represents the latitude argument at the end of the second ignition, ω t represents the argument of perigee of the transfer orbit, represents the eccentric anomaly at the start of the second ignition, t start represents the mission execution time, t scan represents the starting time of the measurement and control arc segment search, u scan represents the starting latitude argument of the measurement and control arc segment search;

[0058] If the mission type is inclination adjustment, then:

[0059]

[0060]

[0061]

[0062]

[0063] Among them, represents the eccentric anomaly at the start of the first ignition, E start represents the eccentric anomaly at the mission execution time;

[0064] If the mission type is combined adjustment, then:

[0065]

[0066]

[0067]

[0068]

[0069] In the above-mentioned online trajectory planning method for a space transfer vehicle for deploying multiple satellites, the latitude and angle of satellite deployment are determined according to the measurement and control conditions of the ground station, so that the satellite deployment process is within the measurement and control range of the ground station, including:

[0070] t scan The average number of orbital elements at the time As the starting orbital root number, with Δt as the step length, continuous orbit recursion is performed to obtain the time t of each step n and orbital elements [a n e n i n Ω n ω n u n ], converted to the position point coordinates [X 84 (n) Y 84 (n) Z 84 (n)]; among them, represents the target mean orbit inclination, represents the right ascension of the target mean ascending node, represents the target's average argument of perigee, a n represents the average semi-major axis of the nth recursion point, e n represents the average eccentricity of the nth recursion point, i n represents the average orbital inclination of the nth recursion point, Ω n represents the average right ascension of the ascending node at the nth recursion point, ω n represents the average argument of perigee of the nth recursive point, u n Indicates the average latitude argument of the nth recursion point;

[0071] The position of the i-th ground station is recorded as [x D(i) y D(i) z D(i) ] T , the ellipsoid plumb vector of the i-th ground station is F i , the line vector between the i-th ground station and the aircraft is n i ,but:

[0072]

[0073] According to F i and n i , determine θ i :

[0074] θ i= arccos(n i ·F i )

[0075] where θ i represents the angle between the vector connecting the i-th ground station and the vehicle and the plumb vector of the ground station ellipsoid, and θ i ∈[0, π];

[0076] If θ i - θ om < 0, then flagD i = 0, determining this position point as the measurement and control end point, and recording the latitude argument of this position point If θ i - θ om ≥ 0, then flagD i = 1, determining this position point as the measurement and control start point, and recording the latitude argument of this position point where θ om represents the ground station observation cone angle, and flagD i represents the measurement and control observable flag.

[0077] In the above online trajectory planning method for a space transfer vehicle in multi-orbit satellite deployment, the satellite deployment planning results include: mission type, mission feasibility, estimated fuel consumption m zy , execution time latitude argument u start , the first ignition duration t1, the second ignition duration t2, the first ignition start point the second ignition start point the measurement and control arc start point and the measurement and control arc end point.

[0078] In the above online trajectory planning method for a space transfer vehicle in multi-orbit satellite deployment, if it is determined that the current remaining fuel does not meet the requirements of the satellite deployment mission, then "target mission feasibility" is transmitted to the ground as infeasible, and the autonomous planning process is terminated.

[0079] The present invention has the following advantages:

[0080] (1) The present invention discloses an online trajectory planning method for a space transfer vehicle in multi-orbit satellite deployment, which can adapt to three different target orbit mission types: altitude adjustment, inclination adjustment, and combined adjustment, and has a wide application range.

[0081] (2) The present invention discloses an online trajectory planning method for a space transfer vehicle in multi-orbit satellite deployment, and the planning algorithm has a small computational amount and can meet the on-board online calculation requirements.

[0082] (3) The present invention discloses an online trajectory planning method for a multi-orbit satellite deploying a space transfer vehicle. The latitude argument of the satellite deployment is determined according to the ground station measurement and control conditions, so that the satellite deployment process can be within the measurement and control range of the ground station. BRIEF DESCRIPTION OF THE DRAWINGS

[0083] Figure 1 is a flowchart of an online trajectory planning method for a multi-orbit satellite deploying a space transfer vehicle in an embodiment of the present invention;

[0084] Figure 2 is a schematic diagram of a measurement and control arc segment in an embodiment of the present invention;

[0085] Figure 3 is a schematic diagram of the change in average semi-major axis in an embodiment of the present invention;

[0086] Figure 4 is a schematic diagram of the change in average eccentricity in an embodiment of the present invention;

[0087] Figure 5 is a schematic diagram of the change in average orbital inclination in an embodiment of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0088] To make the objectives, technical solutions, and advantages of the present invention clearer, the following will further describe in detail the disclosed embodiments of the present invention with reference to the accompanying drawings.

[0089] As Figure 1 , in this embodiment, the online trajectory planning method for a multi-orbit satellite deploying a space transfer vehicle includes:

[0090] Step 1: Receive the satellite deployment task uploaded from the ground.

[0091] Step 2: Determine whether the current remaining fuel meets the requirements of the satellite deployment task.

[0092] In this embodiment, if it is determined that the current remaining fuel meets the requirements of the satellite deployment task, then step 3 is executed; otherwise, "target task feasibility" is transmitted to the ground as infeasible, and the autonomous planning process ends.

[0093] Preferably, the following method can be used to determine whether the current remaining fuel meets the requirements of the satellite deployment task:

[0094] First, calculate the required velocity Δv for inclination adjustment i :

[0095]

[0096] where μ represents the gravitational constant, represents the average semi-major axis of the target orbit, and e obje represents the instantaneous eccentricity of the target orbit, and Δi represents the inclination deviation between the current orbit and the target orbit.

[0097] Secondly, calculate the required velocity Δv for semi-major axis adjustment a :

[0098]

[0099]

[0100]

[0101] Wherein, represents the average eccentricity of the target orbit, r p represents the perigee radius of the transfer orbit, e2 represents the average eccentricity of the transfer orbit, represents the average semi-major axis of the transfer orbit, a p represents the average semi-major axis of the current orbit, e p represents the average eccentricity of the current orbit.

[0102] Furthermore, determine the total velocity increment Δv: Δv = [Δv i Δv a T .

[0103] Furthermore, according to the total velocity increment Δv, calculate the required fuel m zy :

[0104]

[0105] Wherein, m0 represents the current vehicle mass, V e represents the jet velocity of the engine.

[0106] Finally, according to m zy , determine whether the current remaining fuel meets the satellite deployment mission requirements. Among them, if m zy is less than the current remaining fuel, it is determined that the current remaining fuel meets the satellite deployment mission requirements; otherwise, it is determined that the current remaining fuel does not meet the satellite deployment mission requirements.

[0107] Preferably, in this embodiment, the method may further include: using the mean orbital elements [a p e p i p Ω p ω p u0] at the current moment t0 as the initial orbital elements, performing one orbit recurrence to obtain the mean orbital elements [a start at the mission execution moment t start e start i startΩ start ω start u start ]; where, Ω p represents the right ascension of the ascending node of the current orbit on average, ω p represents the argument of perigee of the current orbit on average, and u0 represents the argument of latitude of the current orbit on average; a start represents the semi-major axis on average at the execution time, e start represents the eccentricity on average at the execution time, i start represents the inclination of the orbit on average at the execution time, Ω start represents the right ascension of the ascending node of the current orbit on average at the execution time, ω start represents the argument of perigee of the current orbit on average at the execution time.

[0108] Step 3, determine the mission type according to the current mean orbital elements and the target mean orbital elements.

[0109] In this embodiment, the mission types include but are not limited to: altitude adjustment, inclination adjustment, and combined adjustment. Among them:

[0110] If then determine that the mission type is altitude adjustment.

[0111] If then determine that the mission type is inclination adjustment.

[0112] If and then determine that the mission type is combined adjustment.

[0113] Among them, represents the inclination of the target orbit on average, i p represents the inclination of the current orbit on average, a limit represents the discrimination threshold for altitude adjustment, i limit represents the discrimination threshold for inclination adjustment.

[0114] Step 4, calculate the ignition duration and the argument of latitude in the ignition section according to the determined different mission types, and generate an ignition sequence.

[0115] In this embodiment, the methods for calculating the ignition duration and the argument of latitude in the ignition section are as follows:

[0116] If the mission type is altitude adjustment, then:

[0117]

[0118]

[0119]

[0120]

[0121] wherein, represents the engine fuel consumption per second, t1 represents the duration of the first ignition time, and Δv A represents the speed increment required for altitude adjustment at the first ignition, t2 represents the duration of the second ignition time, and Δv C represents the speed increment required for altitude adjustment at the second ignition, represents the latitude argument at the start time of the first ignition, and u start represents the average latitude argument at the execution time, represents the latitude argument at the start time of the second ignition, n0 represents the current orbital angular velocity, and n t represents the transfer orbit angular velocity.

[0122] If the mission type is inclination adjustment, then:

[0123]

[0124]

[0125] If the mission type is combined adjustment, then:

[0126]

[0127]

[0128]

[0129]

[0130] Step 5: According to the generated ignition sequence, perform satellite deployment planning.

[0131] In this embodiment, according to the generated ignition sequence, the start time of the tracking and control arc segment search and the start latitude argument of the tracking and control arc segment search can be calculated to perform the tracking and control arc segment search; wherein, the start time of the tracking and control arc segment search is the time when the orbit transfer is completed; determine the latitude argument for satellite deployment according to the tracking and control conditions of the ground station to keep the satellite deployment process within the tracking and control range of the ground station.

[0132] Preferably, the calculation process of the start time of the tracking and control arc segment search and the start latitude argument of the tracking and control arc segment search is as follows:

[0133] If the mission type is altitude adjustment, then:

[0134]

[0135]

[0136]

[0137]

[0138] Among them, represents the argument of perigee at the end of the first ignition, e tr represents the eccentricity of the transfer orbit, represents the latitude argument at the end of the first ignition, represents the latitude argument at the end of the second ignition, ω t represents the argument of perigee of the transfer orbit, represents the argument of perigee at the start of the second ignition, t start represents the mission execution time, t scan represents the start time of the tracking and control arc search, u scan represents the starting latitude argument of the tracking and control arc search.

[0139] If the mission type is inclination adjustment, then there are:

[0140]

[0141]

[0142]

[0143]

[0144] Among them, represents the argument of perigee at the start of the first ignition, E start represents the argument of perigee at the mission execution time.

[0145] If the mission type is combined adjustment, then there are:

[0146]

[0147]

[0148]

[0149]

[0150] Preferably, the latitude argument for satellite deployment is determined according to the tracking and control conditions of the ground station, so that the satellite deployment process is within the tracking and control range of the ground station, including:

[0151] Taking the scan mean orbital elements at time t as the initial orbital elements, and performing continuous orbit recursion with Δt as the step size to obtain the time t n and orbital elements [a n e ni n Ω n ω n u n ], the converted position point coordinates are [X 84 (n) Y 84 (n) Z 84 (n)]; where, represents the target mean orbital inclination, represents the target mean right ascension of the ascending node, represents the target mean argument of perigee, a n represents the mean semi-major axis of the nth recurrence point, e n represents the mean eccentricity of the nth recurrence point, i n represents the mean orbital inclination of the nth recurrence point, Ω n represents the mean right ascension of the ascending node of the nth recurrence point, ω n represents the mean argument of perigee of the nth recurrence point, u n represents the mean argument of latitude of the nth recurrence point.

[0152] Denote the position point of the ith ground station as [x D(i) y D(i) z D(i) T , the ellipsoidal plumb vector of the ith ground station is F i , the vector connecting the ith ground station and the aircraft is n i , then:

[0153]

[0154] According to F i and n i , determine θ i :

[0155] θ i = arccos(n i · F i )

[0156] where, θ i represents the angle between the vector connecting the ith ground station and the aircraft and the ellipsoidal plumb vector of the ground station, θ i ∈ [0, π].

[0157] If θ i - θ om < 0, then flagD i = 0, determine this position point as the measurement and control end point, record the argument of latitude of this position point If θ i - θ om ≥ 0, then flagD i ​= 1, determine that this position point is the starting point of measurement and control, and record the latitude argument of this position point where θ om represents the ground station observation cone angle, and flagD i represents the measurement and control observable flag.

[0158] Step 6, download the satellite deployment planning result to the ground. If the ground allows execution, execute the task according to the satellite deployment planning result.

[0159] In this embodiment, as shown in Table 1 below, the satellite deployment planning result includes but is not limited to: mission type, mission feasibility, estimated fuel consumption m zy , execution time latitude argument u start , the first ignition duration t1, the second ignition duration t2, the starting point of the first ignition the starting point of the second ignition the starting point of the measurement and control arc segment (i.e., ) and the end point of the measurement and control arc segment (i.e., ).

[0160] Name Downlink Information Task Type Task Feasibility Estimated Fuel Consumption (kg) Latitude Argument at Execution Time (°) First Ignition Duration (s) Second Ignition Duration (s) First Ignition Starting Point (°) Second Ignition Starting Point (°) Starting Point of TT&C Arc (°) End Point of TT&C Arc (°)

[0161] Table 1, based on the above embodiment of the satellite deployment planning result, is illustrated by an example below.

[0162] Suppose the initial velocity position of the space transfer vehicle in the J2000 system is as shown in Table 2 below:

[0163] Name Value rx (m) -2232291 ry (m) -2606796 rz (m) 5605988 vx (m / s) 4420 vy (m / s) -6307 vz (m / s) -1167

[0164] Table 2, the ground note mission information of the initial velocity position in the J2000 system is as shown in Table 3 below:

[0165] Name Value Target Average Semi-Major Axis (m) 6878140 Target Average Eccentricity 0.001 Target Average Orbit Inclination (°) 61 Target Average Right Ascension of Ascending Node (°) 120 Target Average Argument of Perigee (°) 200

[0166] Table 3, the target mean orbital elements

[0167] The output orbit planning result is shown in Table 4:

[0168]

[0169]

[0170] Table 4, the orbit planning result

[0171] The measurement and control arc segment is as Figure 2 shown, the average semi-major axis change is as Figure 3 shown, the average eccentricity change is as Figure 4 shown, and the average orbital inclination change is as Figure 5 shown.

[0172] Although the present invention has been disclosed above in preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solution of the present invention by using the methods and technical content disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present invention without departing from the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0173] The content not described in detail in the specification of the present invention belongs to the well-known technology of those skilled in the art.

Claims

1. An on-line trajectory planning method for a multi-orbit satellite to deploy a space transfer vehicle, characterized in that, Including: Receiving the satellite deployment mission injected on the ground; Judging whether the current remaining fuel meets the requirements of the satellite deployment mission; If it is determined that the current remaining fuel meets the requirements of the satellite deployment mission, then judge the mission type according to the current mean orbital elements and the target mean orbital elements; among them, the mission type includes: altitude adjustment, inclination adjustment, and combined adjustment; According to the determined different mission types, calculate the ignition duration and the latitude argument of the ignition section, and generate an ignition sequence; According to the generated ignition sequence, carry out satellite deployment planning; Transmit the satellite deployment planning result to the ground. If the ground allows execution, execute the mission according to the satellite deployment planning result.

2. The online trajectory planning method for a multi-orbit satellite deploying a space transfer vehicle according to claim 1, characterized in that, Judging whether the current remaining fuel meets the requirements of the satellite deployment mission includes: The calculated speed Δv required for inclination adjustment i : where μ represents the gravitational constant, represents the mean semi-major axis of the target orbit, and e obj represents the instantaneous eccentricity of the target orbit, and Δi represents the inclination deviation between the current orbit and the target orbit; Calculate the required velocity Δv for semi-major axis adjustment a : Among them, represents the average eccentricity of the target orbit, and r p represents the perigee radius of the transfer orbit, and e2 represents the average eccentricity of the transfer orbit, represents the average semi-major axis of the transfer orbit, and a p represents the average semi-major axis of the current orbit, and e p represents the average eccentricity of the current orbit; Determine the total velocity increment Δv: Δv = [Δv i Δv a T ;​ Based on the total velocity increment Δv, m is calculated zy : where m zy represents the estimated fuel consumption, m0 represents the current aircraft mass, and V e represents the engine jet velocity; According to m zy , determine whether the current remaining fuel meets the requirements of the satellite deployment mission; among them, if m zy is less than the current remaining fuel, it is determined that the current remaining fuel meets the requirements of the satellite deployment mission; otherwise, it does not meet the requirements.

3. The online trajectory planning method for a multi-orbit satellite deploying a space transfer vehicle according to claim 2, wherein Judging the mission type according to the current mean orbital elements and the target mean orbital elements, including: If then determine that the task type is height adjustment; If then determine that the task type is inclination angle adjustment; If and then determine that the task type is combination adjustment; Among them, represents the average orbital inclination of the target orbit, i p represents the average orbital inclination of the current orbit, a limit represents the discrimination threshold for altitude adjustment, i limit represents the discrimination threshold for inclination adjustment.

4. The online trajectory planning method for a multi-orbit satellite deploying a space transfer vehicle according to claim 3, wherein According to the determined different mission types, calculating the ignition duration and the latitude argument of the ignition section, and generating an ignition sequence, including: If the mission type is altitude adjustment, then: Among them, represents the engine fuel consumption per second, t1 represents the duration of the first ignition, and Δv A represents the velocity increment required for altitude adjustment during the first ignition, t2 represents the duration of the second ignition, and Δv C represents the velocity increment required for altitude adjustment during the second ignition, represents the latitude argument at the start of the first ignition, and u start represents the latitude argument at the execution moment, represents the latitude argument at the start of the second ignition, n0 represents the current orbital angular velocity, and n t represents the transfer orbit angular velocity; If the mission type is inclination adjustment, then: If the mission type is combined adjustment, then:

5. The online trajectory planning method for a multi-orbit satellite deploying a space transfer vehicle according to claim 4, characterized in that, Also including: Using the mean orbital elements [a p e p i p Ω p ω p u0] at the current moment t0 as the initial orbital elements, perform an orbital propagation once to obtain the mean orbital elements [a start e start e start i start Ω start ω start u start at the mission execution moment t p ; where Ω p represents the right ascension of the ascending node of the current orbit, ω p represents the argument of perigee of the current orbit, and u0 represents the argument of latitude of the current orbit; a start represents the mean semi-major axis at the execution time, e start represents the mean eccentricity at the execution time, i start represents the mean orbital inclination at the execution time, Ω start represents the mean right ascension of the ascending node at the execution time, ω start represents the mean argument of perigee at the execution time.

6. The online trajectory planning method for a multi-orbit satellite-deployed space transfer vehicle according to claim 5, characterized in that According to the generated ignition sequence, carrying out satellite deployment planning, including: According to the generated ignition sequence, calculate the starting time of the tracking and control arc segment search and the starting latitude argument of the tracking and control arc segment search, and carry out the tracking and control arc segment search; among them, the starting time of the tracking and control arc segment search is the time to complete the orbit transfer; Determine the latitude argument of satellite deployment according to the tracking and control conditions of the ground station, so that the satellite deployment process is within the tracking and control range of the ground station.

7. The online trajectory planning method for a multi-orbit satellite deploying a space transfer vehicle according to claim 6, wherein The calculation process of the starting time of the tracking and control arc segment search and the starting latitude argument of the tracking and control arc segment search is as follows: If the mission type is altitude adjustment, then: Among them, represents the argument of perigee at the end of the first ignition, e tr represents the eccentricity of the transfer orbit, represents the argument of latitude at the end of the first ignition, represents the argument of latitude at the end of the second ignition, ω t represents the argument of perigee of the transfer orbit, represents the argument of perigee at the start of the second ignition, t start represents the mission execution time, t scan represents the start time of the TT&C arc segment search, u scan represents the start argument of latitude of the TT&C arc segment search; If the mission type is inclination adjustment, then: Among them, represents the argument of perigee at the start time of the first ignition, E start represents the argument of perigee at the mission execution time; If the mission type is combined adjustment, then:

8. The online trajectory planning method for a multi-orbit satellite-deployed space transfer vehicle according to claim 7, characterized in that, Determine the latitude argument of satellite deployment according to the tracking and control conditions of the ground station, so that the satellite deployment process is within the tracking and control range of the ground station, including: At time t scan The average orbital elements at a given time As the initial orbital elements, with a step size of Δt, perform continuous orbital recursion to obtain the time t at each step n And the orbital elements [a n e n i n Ω n ω n u n , and convert to the position point coordinates [X 84 (n) Y 84 (n) Z 84 (n)]; where Represents the target average orbital inclination Represents the target average right ascension of the ascending node Represents the target average argument of perigee, a n Represents the average semi-major axis at the nth recursion point, e n Represents the average eccentricity at the nth recursion point, i n Represents the average orbital inclination at the nth recursion point, Ω n Represents the average right ascension of the ascending node at the nth recursion point, ω n Represents the average argument of perigee at the nth recursion point, u n Represents the average argument of latitude at the nth recursion point; Denote the position point of the $i$-th ground station as $[x D(i) y D(i) z D(i) $. T The ellipsoidal plumb vector of the $i$-th ground station is $\mathbf{F}$ i The vector connecting the $i$-th ground station and the aircraft is $\mathbf{n}$ i Then: According to F i and n i , determine θ i : θ i = arccos(n i ·F i ) where, θ i represents the angle between the vector connecting the aircraft and the \(i\)-th ground station and the plumb vector of the ground station ellipsoid, θ i ∈[0, π]; If θ i -θ om <0, then flagD i = 0, determine that this position point is the measurement and control end point, and record the latitude argument of this position point If θ i -θ om ≥0, then flagD i = 1, determine that this position point is the measurement and control start point, and record the latitude argument of this position point Among them, θ om represents the ground station observation cone angle, and flagD i represents the measurement and control observable flag.

9. The online trajectory planning method for a multi-orbit satellite deploying a space transfer vehicle according to claim 8, wherein The satellite deployment planning result includes: mission type, mission feasibility, estimated fuel consumption, latitude argument at the execution time, first ignition duration, second ignition duration, starting latitude argument of the first ignition, starting latitude argument of the second ignition, starting point of the tracking and control arc segment, and ending point of the tracking and control arc segment.

10. The online trajectory planning method for a multi-orbit satellite-deployed space transfer vehicle according to claim 9, characterized in that, If it is determined that the current remaining fuel does not meet the requirements of the satellite deployment mission, then transmit "target mission feasibility" as infeasible to the ground and end the autonomous planning process.

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