A method for designing an earth-to-fire transfer orbit and a related device

Abstract

The present application provides a method and related device for designing a Earth-Mars transfer orbit. Based on a preset number of iterations, the method iteratively processes each orbital parameter through a global optimization algorithm to determine a first minimum velocity increment J from the first total velocity increment J1 of the Mars probe after launch obtained in each iteration. 1min . And determine the Earth-Mars transfer orbit according to the first assignment result of each orbital parameter corresponding to J 1min . The orbital parameters in the embodiments of the present application include the orbit insertion time t0 of the Mars probe, the time dt from t0 to the periselene point EL , the time dt from t0 to the deep space maneuver DSM , the time dt from t0 to Mars EM , the lunar gravity assist radius r PL and the lunar gravity assist azimuth angle #imgabs0#. The Earth-Mars transfer orbit determined based on the above orbital parameters can correct the magnitude and direction of the velocity of the probe when it exits the Earth's sphere of influence through lunar gravity assist, thereby reducing the fuel consumption required to control the Mars probe to enter the predetermined Earth-Mars transfer.

Description

Technical Field

[0001] The present invention relates to the field of deep space exploration technology, and in particular to a method for designing an Earth-Mars transfer orbit and related devices. Background Art

[0002] With the development of science and technology, deep space exploration is one of the most cutting-edge scientific and innovative tasks in the aerospace field. Deep space exploration can not only enhance human's understanding of the unknown areas of the universe and the origin of life, but also promote the development of space science and technology and promote the development and utilization of space resources. Mars, as a neighboring planet of the Earth, has always been one of the most attractive targets for deep space exploration. The first problem to be solved in the design of the orbit of the Mars exploration mission is how to use a launch vehicle to send the probe into the predetermined Earth-Mars transfer orbit.

[0003] Currently, the launch and parking orbits are subject to technical limitations and have a short upper limit on the gliding time. This will cause the orbit that the launch vehicle can reach to be mismatched with the Earth-Mars transfer orbit, and the probe cannot be directly sent into the predetermined Earth-Mars transfer orbit. In related technologies, deep space maneuvers are often used to change the original flight speed and direction of the probe, so that it can smoothly reach Mars along the changed orbit. Since deep space maneuvers will consume the fuel carried by the probe itself, the above method completely uses deep space maneuvers to make up for the lack of carrying capacity, which will greatly affect the service life of the probe after it reaches the mission orbit around Mars. Summary of the invention

[0004] An embodiment of the present application provides a method for designing an Earth-Mars transfer orbit, which uses the force of the moon to change the speed and direction of the probe when it leaves the Earth's impact sphere, thereby reducing the fuel consumption required to change the speed and direction of the probe through deep space maneuvers.

[0005] In a first aspect, an embodiment of the present application provides a method for designing an Earth-Mars transfer trajectory, the method comprising:

[0006] Obtain orbital parameters and preset value ranges for each orbital parameter, wherein the orbital parameters include: the time t0 when the Mars probe enters orbit, the time dt from t0 to the perigee EL , the time from t0 to the deep space maneuver dt DSM , the time from t0 to Mars dt EM , the lunar leverage radius r PL and the lunar leverage azimuth

[0007] Based on the preset number of iterations, each orbital parameter is iteratively processed by the global optimization algorithm to determine the first minimum velocity increment J from the first total velocity increment J1 of the Mars probe after launch obtained in each iteration. 1min ; wherein the J 1mmis the minimum value of J1 obtained in each iteration;

[0008] Determine the J 1min The first assignment results of the corresponding orbital parameters are used to determine the Earth-Mars transfer orbit based on the first assignment results of the orbital parameters; wherein each iteration process is as follows:

[0009] For each orbital parameter, assign a value to the orbital parameter based on a preset value interval of the orbital parameter;

[0010] For each orbital parameter after the assignment, the position vector r of the Mars probe after the lunar leverage is determined based on the orbital integration formula and the leverage flight principle. 1+ and the velocity vector v 1+ ;

[0011] According to the r1+ and the v 1+ Determine the first transfer orbit parameters, and determine the total velocity increment J1 of the Mars probe after launch according to the first transfer orbit parameters, wherein the first transfer orbit parameters include the velocity vector v of the Mars probe after deep space maneuvers 1dsm+ and the velocity vector v when arriving at Mars 1m .

[0012] The embodiment of the present application iterates each orbit parameter through a global optimization algorithm based on a preset number of iterations, so as to determine the first minimum velocity increment J from the first total velocity increment J1 of the Mars probe after launch obtained in each iteration. 1min . And according to J 1min The first assignment result of each corresponding orbital parameter determines the Earth-Mars transfer orbit. The orbital parameters in the embodiment of the present application include the time t0 when the Mars probe enters orbit, the time dt from t0 to the perigee, and the time dt from t0 to the perigee. EL , the time from t0 to the deep space maneuver dt DSM , the time from t0 to Mars dt EM , the lunar leverage radius r PL and the lunar leverage azimuth The Earth-Mars transfer orbit determined based on the above-mentioned orbital parameters can use the force of the moon to correct the speed and direction of the probe when it leaves the Earth's influence sphere, thereby reducing the fuel consumption required to control the Mars probe to enter the predetermined Earth-Mars transfer.

[0013] In some possible embodiments, the position vector r of the Mars probe after leveraging the moon is determined based on the orbital integration formula and the leveraging flight principle. 1+ and the velocity vector v 1+ ,include:

[0014] According to the orbit perigee height r PE , orbital inclination iE And the t0 and dt after assignment EL , taking the orbit transfer angle θ from the entry point to the moon as the iterative variable, determine the initial orbital parameters of the Mars probe rendezvous with the moon, the initial orbital parameters at least include the position vector r of the Mars probe at time t0 01 and the velocity vector v 01 ;

[0015] Based on the orbital integration formula, determining the position vector r1- and velocity vector v1- of the Mars probe entering the lunar influence sphere according to the initial orbital parameters;

[0016] Based on the formula of the principle of leveraging flight, according to the r 1- 、The v 1- 、 PL and stated Determine the r 1+ and the v 1+ .

[0017] After determining the initial orbital parameters of the Mars rover's intersection with the moon, the embodiment of the present application performs orbital integration on the initial orbital parameters to determine the position and speed of the Mars rover when it reaches the lunar sphere of influence, and then determines the position and speed of the Mars rover after leveraging the moon based on the principle formula of leveraging flight.

[0018] In some possible embodiments, the first transfer orbit parameter is determined according to the following method:

[0019] According to the assigned t0 and dt EL Determine the time t1 when the Mars probe arrives at the moon;

[0020] Based on the orbital integration formula, determine the first probe parameter according to the v1+, the r1+ and the t1, the first probe parameter at least including the speed, position and time of the Mars probe entering the Earth's sphere of influence, the first probe parameter being the parameter of the Mars probe in the geocentric coordinate system;

[0021] The first detector parameter is converted based on the cone stitching principle to obtain the second detector parameter, where the second detector parameter is the parameter of the Mars probe in the heliocentric coordinate system;

[0022] Based on the orbital integral formula, according to the second detector parameter and the assigned dt DSM Determine the time t at which the Mars probe arrives at the deep space maneuvering point 1dsm 、Position vector r 1dsm and the velocity vector v 1dsm- , and according to the t 1dsm The dtEM and the query ephemeris JPL after assignment to determine the time t at which the Mars probe arrives at Mars 1m and the Mars rover at t 1m The position vector r at the time 1m ;

[0023] Based on Lambert's formula, according to the t 1dsm 、 1dsm , the 1m And the r 1m , to determine the velocity vector v after deep space maneuvering 1dsm+ and the Mars arrival velocity vector v 1m , thereby determining the first orbit transfer parameters.

[0024] The embodiment of the present application determines the speed, position and time of the Mars probe entering the Earth's sphere of influence based on the orbital integral formula in advance. The above parameters are all parameters of the geocentric coordinate system, and coordinate conversion is required based on the cone splicing principle to obtain the corresponding heliocentric coordinate system parameters. After determining the time when the Mars probe arrives at Mars and its position at that time based on the heliocentric coordinate system parameters, the first orbit transfer parameters are determined based on the Lambert formula.

[0025] In some possible embodiments, determining the total velocity increment J1 of the Mars probe after launch according to the transfer orbit parameters includes:

[0026] According to the v 1dsm- 、The v 1dsm+ and the v 1m Determine the velocity increment Δv required to perform deep space maneuvers on the Mars parametric instrument 1DSM and the velocity increment Δv required for fire capture IMOI ;

[0027] According to the Δv 1DSM and the Δv 1MOI The sum determines the J1.

[0028] In the embodiment of the present application, the sum of the speed increment required for the Mars probe to perform deep space maneuvers and the speed increment required for Mars capture is taken as the total speed increment. The minimum speed increment selected in this way can ensure that the minimum fuel is required for deep space maneuvers.

[0029] In some possible embodiments, determining the Earth-Mars transfer orbit based on the first assignment results of the orbital parameters includes:

[0030] For each orbital parameter, the first value assignment result of the orbital parameter is used as the initial value of the orbital parameter, and each orbital parameter is iteratively processed by the step-size acceleration algorithm to determine the second minimum velocity increment J from the second total velocity increment J2 of the Mars probe after launch obtained in each iteration. 2min ; wherein the J 2min is the minimum value of J2 obtained in each iteration;

[0031] Determine the J 2min The corresponding second assignment results of each orbital parameter are used to determine the Earth-Mars transfer orbit based on the second assignment results of each orbital parameter.

[0032] After determining the first value assignment results of each orbital parameter, the embodiment of the present application uses the first value assignment results of each orbital parameter as the initial value and adopts the step-length acceleration algorithm to perform secondary iterative calculations, thereby accurately determining the values ​​of each parameter and obtaining a more accurate Earth-Mars transfer orbit.

[0033] In some possible embodiments, each iteration process is as follows:

[0034] For each orbital parameter after the assignment, the position vector and velocity vector of the Mars probe in the Earth-Mars transfer process are integrated using an accurate perturbation model to obtain the perigee parameters of the Mars probe; wherein the second transfer orbital parameters include the velocity vector v of the Mars probe after the deep space maneuver 2dsm+ And the velocity vector v when reaching the perigee 2m ;

[0035] The near-fire point parameter is corrected by using a differential correction method so that the corrected near-fire point parameter reaches a preset parameter threshold;

[0036] The second transfer orbit parameters are corrected based on the corrected perigee parameters, and the J2 is determined according to the corrected second transfer orbit parameters.

[0037] In the process of iterating using the step-size acceleration algorithm, the embodiment of the present application takes into account the influence of the perigee parameters on the Earth-Mars transfer orbit, and corrects the second transfer orbit parameters by using the corrected perigee parameters to obtain J2 that satisfies the perigee parameters.

[0038] In some possible embodiments, determining J2 according to the corrected second transfer orbit parameter includes:

[0039] Determine the velocity increment Δv required for the Mars probe to perform deep space maneuvers according to the corrected second transfer orbit parameters 2DSM and the velocity increment Δv required for fire capture 2MOI ;

[0040] According to the Δv 2DSM and the Δv 2MOI The sum determines the J2.

[0041] In the embodiment of the present application, the sum of the speed increment required for the Mars probe to perform deep space maneuvers and the speed increment required for Mars capture is taken as the total speed increment. The minimum speed increment selected in this way can ensure that the minimum fuel is required for deep space maneuvers.

[0042] In a second aspect, an embodiment of the present application provides a device for designing an Earth-Mars transfer trajectory, the device comprising:

[0043] The parameter acquisition module is configured to execute the acquisition of orbital parameters and the preset value range of each orbital parameter, wherein the orbital parameters include: the time t0 when the Mars probe enters orbit, the time dt from t0 to the perigee, and the time dt when the Mars probe enters orbit. EL , the time from t0 to the deep space maneuver dt DSM , the time from t0 to Mars dt EM , the lunar leverage radius r PL and the lunar leverage azimuth

[0044] The velocity increment module is configured to perform iterative processing of each orbit parameter through a global optimization algorithm based on a preset number of iterations, so as to determine a first minimum velocity increment J from the first total velocity increment J1 of the Mars probe after launch obtained in each iteration. 1min ; wherein the J 1min is the minimum value of J1 obtained in each iteration;

[0045] A trajectory generation module is configured to perform the determination of the J 1min The first assignment results of the corresponding orbital parameters are used to determine the Earth-Mars transfer orbit based on the first assignment results of the orbital parameters; wherein each iteration process is as follows:

[0046] For each orbital parameter, assign a value to the orbital parameter based on a preset value interval of the orbital parameter;

[0047] For each orbital parameter after the assignment, the position vector r of the Mars probe after the lunar leverage is determined based on the orbital integration formula and the principle of leveraged flight. 1+ and the velocity vector v 1+ ;

[0048] According to the r1+ and the v 1+ Determine the first transfer orbit parameters, and determine the total velocity increment J1 of the Mars probe after launch according to the first transfer orbit parameters, wherein the first transfer orbit parameters include the velocity vector v of the Mars probe after deep space maneuvers 1dsm+and the velocity vector v when arriving at Mars 1m .

[0049] In some possible embodiments, the method of determining the position vector r of the Mars probe after leveraging the moon based on the orbital integral formula and the leveraging flight principle is performed. 1+ and the velocity vector v 1+ , the track generation module is configured as follows:

[0050] According to the orbit perigee height r PE , orbital inclination i E And the t0 and dt after assignment EL , taking the orbit transfer angle θ from the entry point to the moon as the iterative variable, determine the initial orbit parameters of the Mars probe rendezvous with the moon, the initial orbit parameters at least include the position vector r of the Mars probe at time t0 01 and the velocity vector v 01 ;

[0051] Based on the orbital integration formula, determining the position vector r1- and velocity vector v1- of the Mars probe entering the lunar influence sphere according to the initial orbital parameters;

[0052] Based on the formula of the principle of leveraging flight, according to the r 1- 、The v 1- 、 PL and stated Determine the r 1+ and the v 1+ .

[0053] In some possible embodiments, the first transfer orbit parameter is determined according to the following method:

[0054] According to the assigned t0 and dt EL Determine the time t1 when the Mars probe arrives at the moon;

[0055] Based on the orbital integration formula, determine the first probe parameter according to the v1+, the r1+ and the t1, the first probe parameter at least including the speed, position and time of the Mars probe entering the Earth's sphere of influence, and the first probe parameter is the parameter of the Mars probe in the geocentric coordinate system;

[0056] The first detector parameter is converted based on the cone stitching principle to obtain the second detector parameter, where the second detector parameter is the parameter of the Mars probe in the heliocentric coordinate system;

[0057] Based on the orbital integral formula, according to the second detector parameter and the assigned dt DSMDetermine the time t at which the Mars probe arrives at the deep space maneuvering point 1dsm 、Position vector r 1dsm and the velocity vector v 1dsm- , and according to the t 1dsm The dt EM and the query ephemeris JPL after assignment to determine the time t at which the Mars probe arrives at Mars 1m and the Mars rover at t 1m The position vector r at the time 1m ;

[0058] Based on Lambert's formula, according to the t 1dsm 、 1dsm , the 1m And the r 1m , to determine the velocity vector v after deep space maneuvering 1dsm+ and the Mars arrival velocity vector v 1m , thereby determining the first orbit transfer parameters.

[0059] In some possible embodiments, to determine the total velocity increment J1 of the Mars probe after launch according to the transfer orbit parameters, the orbit generation module is configured as follows:

[0060] According to the v 1dsm- 、The v 1dsm+ and the v 1m Determine the velocity increment Δv required to perform a deep space maneuver on the Mars rover 1DSM and the velocity increment Δv required for Mars capture 1MOI ;

[0061] According to the Δv 1DSM and the Δv 1MOI The sum determines the J1.

[0062] In some possible embodiments, the determining of the Earth-Mars transfer orbit based on the first assignment result of each orbit parameter is performed, and the orbit generation module is configured as follows:

[0063] For each orbital parameter, the first value assignment result of the orbital parameter is used as the initial value of the orbital parameter, and each orbital parameter is iteratively processed by the step-size acceleration algorithm to determine the second minimum velocity increment J from the second total velocity increment J2 of the Mars probe after launch obtained in each iteration. 2min ; wherein the J 2min is the minimum value of J2 obtained in each iteration;

[0064] Determine the J 2minThe corresponding second assignment results of each orbital parameter are used to determine the Earth-Mars transfer orbit based on the second assignment results of each orbital parameter.

[0065] In some possible embodiments, each iteration process is as follows:

[0066] For each orbital parameter after the assignment, the position vector and velocity vector of the Mars probe in the Earth-Mars transfer process are integrated using an accurate perturbation model to obtain the perigee parameters of the Mars probe; wherein the second transfer orbital parameters include the velocity vector v of the Mars probe after the deep space maneuver 2dsm+ and the velocity vector v when arriving at Mars 2m ;

[0067] The near-fire point parameter is corrected by using a differential correction method so that the corrected near-fire point parameter reaches a preset parameter threshold;

[0068] The second transfer orbit parameters are corrected based on the corrected perigee parameters, and the J2 is determined according to the corrected second transfer orbit parameters.

[0069] In some possible embodiments, the determining of J2 according to the corrected second transfer orbit parameter is performed, and the orbit generation module is configured as follows:

[0070] Determine the velocity increment Δv required for the Mars probe to perform deep space maneuvers according to the corrected second transfer orbit parameters 2DSM and the velocity increment Δv required for fire capture 2MOI ;

[0071] According to the Δv 2DSM and the Δv 2MOI The sum determines the J2.

[0072] In a third aspect, an embodiment of the present application further provides an electronic device, including:

[0073] processor;

[0074] a memory for storing instructions executable by the processor;

[0075] The processor is configured to execute the instructions to implement any one of the methods provided in the first aspect of the present application.

[0076] In a fourth aspect, an embodiment of the present application further provides a computer-readable storage medium. When the instructions in the computer-readable storage medium are executed by a processor of an electronic device, the electronic device is enabled to execute any of the methods provided in the first aspect of the present application.

[0077] Other features and advantages of the present application will be described in the following description, and partly become apparent from the description, or be understood by practicing the present application. The purpose and other advantages of the present application can be realized and obtained by the structures specifically pointed out in the written description, claims, and drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0078] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required for use in the embodiments of the present application will be briefly introduced below. Obviously, the drawings introduced below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0079] Figure 1 A schematic diagram of the Earth-Mars transfer orbit shown in an embodiment of the present application;

[0080] Figure 2 This is an overall flow chart of a method for designing an Earth-Mars transfer trajectory shown in an embodiment of the present application;

[0081] Figure 3 A flowchart of determining a first value assignment result of a track parameter shown in an embodiment of the present application;

[0082] Figure 4 A flow chart of determining a second value assignment result of a local optimization parameter shown in an embodiment of the present application;

[0083] Figure 5 This is a structural diagram of an Earth-Mars transfer orbit design device 500 shown in an embodiment of the present application;

[0084] Figure 6 This is a schematic diagram of an electronic device shown in an embodiment of the present application. DETAILED DESCRIPTION

[0085] The technical solutions in the embodiments of the present application will be described clearly and in detail below in conjunction with the accompanying drawings. In the description of the embodiments of the present application, unless otherwise specified, " / " will mean or, for example, A / B can mean A or B; "and / or" in the text is only a description of the association relationship of associated objects, indicating that there can be three relationships, for example, A and / or B can mean: A exists alone, A and B exist at the same time, and B exists alone. In addition, in the description of the embodiments of the present application, "multiple" means two or more than two.

[0086] In the description of the embodiments of the present application, unless otherwise specified, the term "multiple" refers to two or more, and other quantifiers are similar and should be understood. The preferred embodiments described herein are only used to illustrate and explain the present application, and are not used to limit the present application. In addition, the embodiments of the present application and the features therein may be combined with each other if there is no conflict.

[0087] To further illustrate the technical solution provided by the embodiment of the present application, this is described in detail below in conjunction with the accompanying drawings and specific implementation methods. Although the embodiment of the present application provides the method operation steps as shown in the following embodiments or drawings, more or fewer operation steps may be included in the method based on routine or no creative labor. In the steps where there is no necessary causal relationship logically, the execution order of these steps is not limited to the execution order provided by the embodiment of the present application. The method can be executed in the order of the method shown in the embodiment or drawings or in parallel during the actual processing process or when the control device is executed.

[0088] As mentioned above, the first problem to be solved in the orbit design of the Mars exploration mission is how to use the launch vehicle to send the probe into the predetermined Earth-Mars transfer orbit. From the perspective of the orbit, if you want to design an Earth-Mars transfer orbit that can meet the requirements of the space mission and satisfy various transportation constraints, the transportation constraints are mainly divided into two parts. One is the energy constraint of the entry orbit, and the other is the direction of the entry orbit in space. The energy constraint can be met by adjusting the time of launch from the Earth and arrival at Mars, that is, setting a suitable exploration cycle (for example, the Mars exploration cycle is about 26 months), and the direction of the entry orbit in space is related to multiple factors such as the location of the launch site, the launch direction, and the length of the glide segment. At present, the launch parking orbit is subject to technical limitations and has the problem of low glide time, resulting in the mismatch between the entry orbit that the launch vehicle can reach and the Earth-Mars transfer orbit, and the probe cannot be directly sent into the predetermined Earth-Mars transfer orbit.

[0089] In related technologies, deep space maneuvers are often used to change the original flight speed and direction of the probe, so that it can smoothly enter the Earth-Mars transfer orbit along the changed orbit. Since deep space maneuvers will consume the fuel carried by the probe itself, if deep space maneuvers are used entirely to make up for the lack of carrying capacity, it will greatly affect the life of the probe after it reaches the mission orbit around Mars.

[0090] To solve the above problems, the invention of the present application is as follows: based on a preset number of iterations, each orbital parameter is iteratively processed by a global optimization algorithm to determine a first minimum velocity increment J from the first total velocity increment J1 of the Mars probe after launch obtained in each iteration. 1min . And according to J 1minThe first assignment result of each corresponding orbital parameter determines the Earth-Mars transfer orbit. The orbital parameters in the embodiment of the present application include the time t0 when the Mars probe enters orbit, the time dt from t0 to the perigee, and the time dt from t0 to the perigee. EL , the time from t0 to the deep space maneuver dt DSM , the time from t0 to Mars dt EM , the lunar leverage radius r PL and the lunar leverage azimuth The Earth-Mars transfer orbit determined based on the above orbital parameters can use the gravity of the moon to assist in correcting the speed and direction of the probe when it leaves the Earth's influence sphere, thereby reducing the fuel consumption required to change the speed and direction of the probe through deep space maneuvers.

[0091] Figure 1 The schematic diagram of the Earth-Mars transfer orbit provided in the embodiment of the present application can be specifically as follows Figure 1 As shown, the technical solution provided in the embodiment of the present application can control the Mars rover to fly around the moon, so as to correct the speed and direction of the rover when it leaves the Earth's sphere of influence by leveraging the force of the moon. That is, the speed and direction of the Mars rover leaving the Earth's sphere of influence are corrected by leveraging the force of the moon. Although the ability to correct the speed and direction of the rover in this way is limited, it is ultimately necessary to change the speed and direction of the rover through deep space maneuvers so that the Mars rover can smoothly reach Mars along the changed orbit. However, leveraging the force of the moon will greatly reduce the fuel consumption required for deep space braking, thereby increasing the service life of the Mars rover after it reaches the mission orbit around Mars.

[0092] To facilitate understanding of a method for designing an Earth-Mars transfer orbit provided in an embodiment of the present application, Figure 2 As shown, Figure 2 The overall flow chart of the technical solution of this application includes:

[0093] Step 201: Obtain orbital parameters and a preset value range for each orbital parameter, wherein the orbital parameters include: the time t0 when the Mars probe enters orbit, the time dt from t0 to the perigee EL , the time from t0 to the deep space maneuver dt DSM , the time from t0 to Mars dt EM , the lunar leverage radius r PL and the lunar leverage azimuth

[0094] Step 202: Based on a preset number of iterations, the orbital parameters are iteratively processed by a global optimization algorithm to determine a first minimum velocity increment J from the first total velocity increment J1 of the Mars probe after launch obtained in each iteration. 1min ; wherein the J 1min is the minimum value of J1 obtained in each iteration;

[0095] Step 203: Determine the J 1min The first assignment results of the corresponding orbital parameters are used to determine the Earth-Mars transfer orbit based on the first assignment results of the orbital parameters; wherein each iteration process is as follows:

[0096] For each orbital parameter, assign a value to the orbital parameter based on a preset value interval of the orbital parameter;

[0097] For each orbital parameter after the assignment, the position vector r of the Mars probe after the lunar leverage is determined based on the orbital integration formula and the principle of leveraged flight. 1+ and the velocity vector v 1+ ;

[0098] According to the r1+ and the v 1+ Determine the first transfer orbit parameter, and determine the total velocity increment J1 of the Mars probe after launch according to the first transfer orbit parameter, wherein the first transfer orbit parameter includes the velocity v of the Mars probe after deep space maneuvering 1dsm+ and the speed v when arriving at Mars 1m .

[0099] During implementation, the Earth-Mars transfer orbit constraints and launch constraints are first determined from the engineering mission requirements. The Earth-Mars transfer orbit constraints include the perigee height r of the entry orbit. pE , orbital inclination i E , Near fire point height r pM 、Mars equatorial orbit inclination i M and the height of the far-fire point in the orbit after near-fire capture r aM ; Launch constraints include the maximum orbital energy that the launch can provide and the range of perigee angles (ω min ,ω max ). Then the global optimization algorithm is used to iteratively process the orbital parameters.

[0100] Before the global optimization algorithm is used to iterate each orbital parameter, the global optimization population size and the upper limit of the number of iterations need to be set in advance. Then, the orbital parameters are iteratively processed according to the set global optimization population size and the preset first number of iterations. Each iteration process is as follows:

[0101] For each orbital parameter, a value is assigned to the orbital parameter based on the preset value range of the orbital parameter. It should also be understood that each of the above-mentioned orbital parameters is set within a certain range in the engineering mission. For example, the time t0 when the Mars probe enters orbit may be within a range of 2 to 3 months in the engineering mission. The present application is intended to select the value used to determine the Earth-Mars transfer orbit from the preset value range of each orbital parameter in the engineering mission, so as to determine the Earth-Mars transfer orbit that can correct the speed and direction of the Mars probe when it leaves the Earth's influence sphere through the force of the moon, thereby reducing the fuel consumption required to control the Mars probe to enter the predetermined Earth-Mars transfer.

[0102] It should also be noted that the global optimization algorithm (particle swarm optimization method PSO) will assign values ​​to each orbital parameter based on the preset value range of each orbital parameter. After each assignment of all orbital parameters is completed, a first total velocity increment J1 will be calculated using the assigned orbital parameters.

[0103] For each orbital parameter after the assignment, it is necessary to first determine the position r1+ and speed v1+ of the Mars probe after the lunar leverage based on the orbital integral formula and the principle of leveraged flight. pE , orbital inclination i E And t0 and dt after assignment EL , taking the orbit transfer angle θ from the entry point to the moon as the iteration variable, determine the initial orbital parameters of the Mars probe and the moon rendezvous. The initial orbital parameters at least include the position vector r of the Mars probe at time t0 01 and the velocity vector v 01 .

[0104] The specific iterative solution process includes the following steps S1 to S7:

[0105] Step S1, determine the time t when the Mars probe arrives at the moon 1L Among them, t 1L t0 and dt after assignment EL sum.

[0106] Step S2, determine the moon at t according to the ephemeris (Jet Propulsion Laboratory) 1L Right Ascension λ at the moment L and declination

[0107] Step S3, determining the right ascension Ω of the orbit ascending node and the angular distance u of the ascending node.

[0108] When executing step S3, it is necessary to determine whether the track is in the ascending track intersection or the descending track intersection. When the ascending track intersects, the following formula (1) can be used to determine Ω and u. When the descending track intersects, the following formula (2) can be used to determine Ω and u.

[0109]

[0110]

[0111] Among them, λ L For the moon at t 1L Right ascension at the moment, For the moon at t 1L The declination at the moment, i E is the orbital inclination in the Earth-Mars orbit constraint.

[0112] Step S4, determining the perigee argument ω according to the ascending node angular distance and the orbital transfer angle from the entry point to the moon.

[0113] Among them, ω is the difference between the angular distance of the ascending node and the orbital transfer angle from the entry point to the moon, that is, ω=u-θ.

[0114] Step S5, determine the orbit semi-major axis α and eccentricity e.

[0115] When executing step S5, it is necessary to determine in advance whether the orbit entry point is the perigee according to the engineering task constraints. If it is the perigee, α and e are determined according to the following formula (3); if it is not the perigee, α and e can be determined based on the Lambert equation.

[0116]

[0117] Step S6, determining the transfer time deviation δdt.

[0118] Among them, the transfer time dt and the assigned dt EL The difference is δdt. The transfer time dt can be determined according to the following formula (4):

[0119]

[0120] Among them, n is the average angular velocity of the orbit, μ is the gravitational constant of the Earth, f is the true anomaly of the orbit, E is the eccentric anomaly of the orbit, and M is the mean anomaly of the orbit.

[0121] Step S7, determine whether the value of δdt is within the preset error range. If it is within the preset error range, exit this process, and the initial orbit parameters determined by this process can be used for subsequent calculations. If it is not within the preset error range, it is necessary to iterate the above steps S1 to S6 again until δdt within the preset error range is obtained, and the initial orbit parameters corresponding to the δdt are obtained.

[0122] After the initial orbital parameters are determined through the above process, the position vector r1- and velocity vector v1- of the Mars probe entering the lunar impact sphere need to be determined based on the orbital integration formula according to the initial orbital parameters. And based on the formula of the principle of leveraged flight, according to r1-, v1-, r PL and Determine r 1+ and v 1+ Specifically, according to the moon's geocentric velocity vector v L Determine the velocity vector v1 of the probe relative to the moon when it enters the lunar impact sphere ∞- . And according to v L When the Mars probe leaves the lunar impact sphere, the velocity vector v relative to the moon is ∞+ The sum determines v 1+ .

[0123] v ∞- is the velocity vector of the probe relative to the moon when it enters the lunar influence sphere, obtained by the following formula (5):

[0124] v ∞- =v 1- -v L (5)

[0125] Among them, v L The velocity vector of the moon relative to the earth. δ is v ∞- and v ∞+ The speed angle δ is obtained by the following formula (6):

[0126]

[0127] Among them, μ L is the gravitational constant of the moon, r PL is the leverage radius.

[0128] The above v ∞+ is the speed of the Mars probe out of the lunar impact ball in the leverage coordinate system. It can be determined by the following formula (7):

[0129]

[0130] Among them, i, j, and k are the unit vectors of the three directions of the force coordinate system.

[0131] The above v 1+ is the velocity of the Mars probe relative to the Earth when it leaves the lunar impact ball, which is determined by the following formula (8):

[0132] v 1+ =v L +v ∞+ (8)

[0133] Determine r through the above process 1+ and v 1+ After that, according to the assigned t0 and dt EL Determine the time t1 when the Mars probe arrives at the moon. Then, based on the orbital integration formula, according to v 1+ 、r 1+ And t1 determines the first probe parameter. The first probe parameter in the embodiment of the present application at least includes the speed, position and time of the Mars probe entering the Earth's sphere of influence.

[0134] It should be understood that the above-mentioned first detector parameters are the parameters of the Mars probe in the geocentric coordinate system. Therefore, it is necessary to convert the first detector parameters based on the cone splicing principle to obtain the second detector parameters. The second detector parameters are the parameters of the Mars probe in the heliocentric coordinate system. After the coordinate conversion, based on the orbital integral formula, according to the second detector parameters and the assigned dt DSM Determine the position vector r of the Mars probe when it reaches the deep space maneuvering point 1dsm and the velocity vector v 1dsm- . And according to r 1dsm 、v dsm- And dt after assignment SM Query the ephemeris table JPL to determine the time t when the Mars probe arrives at Mars 1m and the Mars Rover at t m The position at the moment r 1m Among them, t 1m That is, t0 and dt after assignment EM Further, based on the Lambert formula, according to the above r 1dsm ,t 1m and r 1m Determine v 1dsm+ and v 1m .

[0135] In determining v 1dsm+ and v 1m (i.e. the first transfer orbit parameter), it is necessary to calculate the value of v 1dsm+ and v 1m Determine the velocity increment Δv required to perform deep space maneuvers on the Mars rover 1DSM and the velocity increment Δv required for Mars capture 1MOI , then Δv 1DSM With Δv 1MOI The sum is taken as the J1 corresponding to each orbital parameter assigned in this round. 1DSM and Δv 1MOI It is determined according to the following formula (9):

[0136]

[0137] Among them, μ M is the Martian gravitational constant, r 1m is the radius of the Mars sphere of influence MSOI, r PM is the perigee height of Mars capture orbit, r aM The altitude of the apogee of Mars capture orbit.

[0138] Iterate the above process to determine the first total velocity increment J1 corresponding to each orbit parameter after different assignments. Finally, use PSO to solve the optimized orbit, assuming that each orbit parameter is x i , i∈[1,6] and i is a positive integer. Determine J according to the following formula (10): 1min The first assignment result of the corresponding orbital parameters.

[0139]

[0140] Among them, [t 0min , t 0max ] is the launch window interval of the Mars probe; [dt ELmin , dt ELmax ] is the time range of Earth-Moon transfer, [dt EMmin , dt EMmax ] is the total time range of the Earth-Mars transfer; (0, dt EM ) is the deep space maneuvering time range; [r PLmin , r PLmax ] is the range of the leverage radius, r PLmin It must be larger than the radius of the moon and have a safety margin; is the value range of the leverage azimuth angle.

[0141] The Earth-Mars transfer orbit determined based on the orbital parameters in the above process can use the force of the moon to correct the speed and direction of the probe when it leaves the Earth's influence sphere, thereby reducing the fuel consumption required to change the speed and direction of the probe through deep space maneuvers. However, considering that in actual engineering tasks, it is necessary to consider the impact of perigee-related parameters on the orbit, this application uses a more accurate step-size acceleration algorithm to locally optimize the orbit after determining the globally optimized orbital parameters through the above process.

[0142] It should also be noted that after a large number of global optimization tests, it was found that when the first total velocity increment J1 is the smallest, the lunar leverage radius r PL Usually the minimum value is taken. In order to save computing power, in the subsequent local optimization process of each orbit parameter, the embodiment of the present application will fix r PL The minimum value of r PLThe first assignment result of is taken as the final solution of the orbital parameters. In addition, the embodiment of the present application also replaces dt in the above orbital parameters with the initial orbital semi-major axis α which is more suitable for the accurate perturbation model. EL , and using the perigee orbit inclination i L Substitute the above orbital parameters The Earth-Mars transfer orbit established based on the locally optimized orbital parameters takes into account the impact of perigee-related parameters on the orbit and meets the orbit requirements of actual engineering tasks.

[0143] For the above orbital parameters (i.e. t0, α, i L 、dt EM and dt DSM ) is used for local optimization, the first value assignment result of the orbital parameter is used as the initial value of the orbital parameter in advance, and each orbital parameter is iteratively processed by the step-length acceleration algorithm to determine the second minimum velocity increment J from the second total velocity increment J2 of the Mars probe after launch obtained in each iteration. 2min Among them, J 2min is the minimum value of J2 obtained in each iteration. Then determine J 2min The second value assignment results of the corresponding orbital parameters are used to determine the Earth-Mars transfer orbit. To facilitate understanding of the technical solution of the embodiment of the present application, the above orbital parameters are replaced by local optimization parameters in the following text.

[0144] Each iteration process using the step acceleration method is as follows:

[0145] For each local optimization parameter after assignment, the position and velocity of the Mars probe during the Earth-Mars transfer process are integrated using an accurate perturbation model to obtain the perigee parameters of the Mars probe; among them, the second transfer orbit parameters include the velocity vector v of the Mars probe after deep space maneuvers 2dsm+ And the velocity vector v when reaching the perigee 2m .

[0146] The differential correction method is used to correct the perigee parameters so that the corrected perigee parameters reach the preset parameter threshold. Then, the second transfer orbit parameters are corrected based on the corrected perigee parameters, and J2 is determined based on the corrected second transfer orbit parameters. The perigee parameters in the embodiment of the present application include the perigee time t moi , the height of the fire point r pm , perigee orbit inclination i pm .

[0147] During implementation, the deep space maneuvering speed increment Δv based on the global optimization in the previous text 1dsm =(Δv x , Δv y , Δv z) T As the initial value, use the differential correction method to correct Δv 1dsm Correction is performed to obtain the corrected Δv 2dsm , so that the near-fire point parameters meet the requirements of the engineering task, and then obtain the locally optimized target parameter values.

[0148] The specific process of differential correction is to determine the velocity vector v before deep space maneuver according to the local optimization parameters. 1dsm- , v 1dsm- With Δv 1dsm The sum is the velocity vector v after deep space maneuvering 1dsm+ Then based on the target near-fire parameter variable q = (r pm ,i pm ) T v 1dsm+ The correction relationship is shown in the following formula (11):

[0149] q=f(v 1dsm+ ) (11)

[0150] It should be noted that the deviation of the above target variable value relative to the specified target variable value is Δq=qq * , (q * is the preset expectation, is a known value). After determining the deviation Δq, the speed after deep space maneuver correction is calculated according to the following formula (12).

[0151] v 2dsm+ =E -1 Δq (12)

[0152] Among them, the matrix E -1 State transfer matrix. The velocity of the correction point on the Mars rover's Earth-Mars transfer path is updated through the above process, and the loop is iterated until Δq meets the convergence requirements.

[0153] In some possible embodiments, determining J2 according to the modified second transfer local optimization parameter includes:

[0154] Determine the velocity increment Δv required for deep space maneuvers of the Mars probe based on the modified second transfer local optimization parameters 2DSM and the velocity increment Δv required for fire capture 2MOI , according to Δv 2DSM and Δv 2MOI The sum determines J2.

[0155] During implementation, the J2 obtained by the first assignment of local optimization parameters is used as the target value, and the local optimization parameters are assigned a second time with a preset step size. If the J2 obtained by the local optimization parameters after the second assignment meets the preset convergence requirements, the values ​​of the local optimization parameters after the second assignment are used as the second assignment results of the local optimization parameters. 2’ If the preset convergence requirement is not met, the local optimization parameters are assigned a third time based on the preset step size until the minimum speed increment corresponding to the local optimization parameters after the assignment meets the convergence requirement.

[0156] When assigning values, the first assignment result of each local optimization parameter shall be used as the initial value of the first assignment. When assigning values ​​again, the preset step size shall be added or subtracted based on the first assignment. In the embodiment of the present application, in order to obtain the minimum speed increment of the Mars rover after deep hole maneuvering, if the minimum speed increment J corresponding to each local optimization parameter assigned for the nth time is 2(n) The minimum speed increment J corresponding to each local optimization parameter assigned less than the n-1th time 2(n-1) , then in the n+1th assignment, the preset step length needs to be added to the nth assignment; otherwise, the preset step length needs to be reduced based on the nth assignment until the minimum speed increment corresponding to each local optimization parameter after the assignment meets the convergence requirement.

[0157] After obtaining the second assignment results of each local optimization parameter through the above process, the whole process can be verified through simulation. The optimal solution obtained by local optimization is used as the initial orbit. The orbital parameters of the entire Earth-Mars transfer process are calculated under the precise perturbation model to verify the initial parameters of the Earth departure and the terminal perigee parameters, as well as the correctness of the transfer process.

[0158] To facilitate understanding of how to determine the first value assignment result of each orbital parameter in the above process, the specific steps are as follows: Figure 3 As shown, the following steps are included:

[0159] Step 301: Determine the population size and preset number of iterations of the global optimization algorithm;

[0160] Step 302: Obtain 6 orbital parameters; each orbital parameter is the time t0 when the Mars probe enters orbit, the time dt from t0 to the perigee EL , the time from the Earth's impact ball to the deep space maneuver dt DSM , the time from the Earth's impact on the sphere to the Mars' impact on the sphere dt EM , the lunar leverage radius r PL and the lunar leverage azimuth

[0161] Step 303: using a global optimization algorithm to assign values ​​to each orbital parameter, taking the orbital transfer angle from the entry point to the moon as an iterative variable, iteratively solving the orbit of the Mars probe rendezvous with the moon, and then obtaining the initial orbital parameters;

[0162] Step 304: Determine whether the orbit energy and perigee argument of the orbital entry meet the launch constraints;

[0163] Step 305: If the constraints are met, orbital integration is performed based on the initial orbital parameters to obtain the position velocity of the lunar impact sphere. Based on the principle formula of leveraged flight, the position velocity of the lunar impact sphere after leverage can be obtained. If not, return to step 303 and re-assign each orbital parameter until the launch constraints are met.

[0164] Step 306: Use the position and velocity of the probe after the lunar gravity assist, orbitally integrate to the Earth's influence sphere, and use the cone splicing principle to obtain the initial position and velocity of the probe eye center in the influence sphere. Then extrapolate the heliocentric orbit to the deep space maneuvering point to obtain the position and velocity of the deep space maneuvering point;

[0165] Step 307: Obtain the time when the Mars probe arrives at Mars and the position of Mars at that time from the deep space maneuvering position and the JPL ephemeris table;

[0166] Step 308: Use the Lambert equation to solve the transfer orbit from deep space maneuver to Mars, and obtain the velocity of the probe after deep space maneuver and the velocity at Mars;

[0167] Step 309: Determine the speed increment required for deep space maneuvering and the speed increment required for Mars capture by using the speed of the probe after deep space maneuvering and the speed at Mars, thereby determining the first total speed increment J1 of the Mars probe after launch. Wherein, J1 is the sum of the speed increment required for deep space maneuvering and the speed increment required for Mars capture;

[0168] Step 310: Determine whether the current number of iterations reaches the upper limit N;

[0169] Step 311: When the upper limit N is reached, the first minimum speed increment J is determined from J1 obtained in each iteration. 1min , J 1min The value of each corresponding orbital parameter is the first assignment result of the orbital parameter. If the upper limit N is not reached, return to step 303, re-assign each orbital parameter, and determine the J1 corresponding to each orbital parameter after the assignment through steps 303 to 309.

[0170] To facilitate understanding of how to determine the second assignment result of each local optimization parameter in the above process, the specific method can be as follows: Figure 4 As shown, the following steps are included:

[0171] Step 401: determine local optimization parameters, and use a step-length acceleration algorithm to take the first assignment result of each local optimization parameter as an initial value;

[0172] Step 402: using an accurate perturbation model to integrate the position vector and velocity vector of the Mars probe during the Earth-Mars transfer process to obtain perigee parameters of the Mars probe;

[0173] Step 403: Integrate to the near-fire point using the accurate perturbation model to obtain near-fire point parameters;

[0174] Step 404: determining whether the near-fire point parameters meet the convergence criteria;

[0175] Step 405: If the convergence standard is not met, the deep space maneuver control amount is corrected by using a differential correction method so that the perigee parameter meets the convergence standard;

[0176] Step 406: If the convergence criterion is met, further determining whether the local optimization parameters have converged;

[0177] Step 407: if convergence has not yet occurred, re-assign values ​​to each local optimization parameter using a preset step size, and after the assignment, return to step 402 to execute the above process of steps 402 to 407;

[0178] Step 408: Convergence, determine J2 as J 2min , and determine J 2min The second assignment result corresponding to each local optimization parameter.

[0179] Based on the same inventive concept, the present application embodiment provides a device 500 for designing an Earth-Mars transfer trajectory, specifically, Figure 5 As shown, including:

[0180] The parameter acquisition module 501 is configured to execute the acquisition of orbital parameters and the preset value range of each orbital parameter, wherein the orbital parameters include: the time t0 when the Mars probe enters orbit, the time dt from t0 to the perigee, and the time dt EL , the time from t0 to the deep space maneuver dt DSM , the time from t0 to Mars dt EM , the lunar leverage radius r PL and the lunar leverage azimuth

[0181] The velocity increment module 502 is configured to perform iterative processing on each orbit parameter through a global optimization algorithm based on a preset number of iterations, so as to determine a first minimum velocity increment J from the first total velocity increment J1 of the Mars probe after launch obtained in each iteration. 1min ; wherein the J 1min is the minimum value of J1 obtained in each iteration;

[0182] The track generation module 503 is configured to perform the determination of the J 1min The first assignment results of the corresponding orbital parameters are used to determine the Earth-Mars transfer orbit based on the first assignment results of the orbital parameters; wherein each iteration process is as follows:

[0183] For each orbital parameter, assign a value to the orbital parameter based on a preset value interval of the orbital parameter;

[0184] For each orbital parameter after the assignment, the position vector r of the Mars probe after the lunar leverage is determined based on the orbital integration formula and the leverage flight principle. 1+ and the velocity vector v 1+ ;

[0185] According to the r1+ and the v 1+ Determine the first transfer orbit parameters, and determine the total velocity increment J1 of the Mars probe after launch according to the first transfer orbit parameters, wherein the first transfer orbit parameters include the velocity vector v of the Mars probe after deep space maneuvers 1dsm+ and the velocity vector v when arriving at Mars 1m .

[0186] In some possible embodiments, the method of determining the position vector r of the Mars probe after leveraging the moon based on the orbital integral formula and the leveraging flight principle is performed. 1+ and the velocity vector v 1+ , the track generation module 503 is configured as follows:

[0187] According to the orbit perigee height r PE , orbital inclination i E And the t0 and dt after assignment EL , taking the orbit transfer angle θ from the entry point to the moon as the iterative variable, determine the initial orbit parameters of the Mars probe rendezvous with the moon, the initial orbit parameters at least include the position vector r of the Mars probe at time t0 01 and the velocity vector v 01 ;

[0188] Based on the orbital integration formula, determining the position vector r1- and velocity vector v1- of the Mars probe entering the lunar influence sphere according to the initial orbital parameters;

[0189] Based on the formula of the principle of leveraging flight, according to the r 1- 、The v 1- 、 PL and stated Determine the r 1+ and the v 1+ .

[0190] In some possible embodiments, the first transfer orbit parameter is determined according to the following method:

[0191] According to the assigned t0 and dt EL Determine the time t1 when the Mars probe arrives at the moon;

[0192] Based on the orbital integration formula, determine the first probe parameter according to the v1+, the r1+ and the t1, the first probe parameter at least including the speed, position and time of the Mars probe entering the Earth's sphere of influence, and the first probe parameter is the parameter of the Mars probe in the geocentric coordinate system;

[0193] The first detector parameter is converted based on the cone stitching principle to obtain the second detector parameter, where the second detector parameter is the parameter of the Mars probe in the heliocentric coordinate system;

[0194] Based on the orbital integral formula, according to the second detector parameter and the assigned dt DSM Determine the time t at which the Mars probe arrives at the deep space maneuvering point 1dsm 、Position vector r 1dsm and the velocity vector v 1dsm- , and according to the t 1dsm The dt EM and the query ephemeris JPL after assignment to determine the time t at which the Mars probe arrives at Mars 1m and the Mars rover at t 1m The position vector r at the time 1m ;

[0195] Based on Lambert's formula, according to the t 1dsm 、 1dsm , the 1m And the r 1m , to determine the velocity vector v after deep space maneuvering 1dsm+ and the Mars arrival velocity vector v 1m , thereby determining the first orbit transfer parameters.

[0196] In some possible embodiments, to determine the total velocity increment J1 of the Mars probe after launch according to the transfer orbit parameters, the orbit generation module 503 is configured as follows:

[0197] According to the v 1dsm- 、The v 1dsm+ and the v 1m Determine the velocity increment Δv required to perform a deep space maneuver on the Mars rover 1DSM and the velocity increment Δv required for Mars capture1MOI ;

[0198] According to the ΔΔv 1DSM and the Δv 1MOI The sum determines the J1.

[0199] In some possible embodiments, the determining of the Earth-Mars transfer orbit based on the first assignment result of each orbit parameter is performed, and the orbit generation module 503 is configured as follows:

[0200] For each orbital parameter, the first value assignment result of the orbital parameter is used as the initial value of the orbital parameter, and each orbital parameter is iteratively processed by the step-size acceleration algorithm to determine the second minimum velocity increment J from the second total velocity increment J2 of the Mars probe after launch obtained in each iteration. 2min ; wherein the J 2min is the minimum value of J2 obtained in each iteration;

[0201] Determine the J 2min The corresponding second assignment results of each orbital parameter are used to determine the Earth-Mars transfer orbit based on the second assignment results of each orbital parameter.

[0202] In some possible embodiments, each iteration process is as follows:

[0203] For each orbital parameter after the assignment, the position vector and velocity vector of the Mars probe in the Earth-Mars transfer process are integrated using an accurate perturbation model to obtain the perigee parameters of the Mars probe; wherein the second transfer orbital parameters include the velocity vector v of the Mars probe after the deep space maneuver 2dsm+ and the velocity vector v when arriving at Mars 2m ;

[0204] The near-fire point parameter is corrected by using a differential correction method so that the corrected near-fire point parameter reaches a preset parameter threshold;

[0205] The second transfer orbit parameters are corrected based on the corrected perigee parameters, and the J2 is determined according to the corrected second transfer orbit parameters.

[0206] In some possible embodiments, to determine J2 according to the modified second transfer orbit parameter, the orbit generation module 503 is configured as follows:

[0207] Determine the velocity increment Δv required for the Mars probe to perform deep space maneuvers according to the corrected second transfer orbit parameters 2DSM and the velocity increment Δv required for fire capture 2MOI ;

[0208] According to the Δv 2DSM and the Δv 2MOI The sum determines the J2.

[0209] Refer to the following Figure 6 The electronic device 130 according to this embodiment of the present application is described. Figure 6 The electronic device 130 shown is merely an example and should not bring any limitation to the functions and scope of use of the embodiments of the present application.

[0210] like Figure 6 As shown, the electronic device 130 is in the form of a general electronic device. The components of the electronic device 130 may include but are not limited to: the at least one processor 131, the at least one memory 132, and a bus 133 connecting different system components (including the memory 132 and the processor 131).

[0211] Bus 133 represents one or more of several types of bus structures, including a memory bus or memory controller, a peripheral bus, a processor, or a local bus using any of a variety of bus architectures.

[0212] The memory 132 may include a readable medium in the form of a volatile memory, such as a random access memory (RAM) 1321 and / or a cache memory 1322 , and may further include a read-only memory (ROM) 1323 .

[0213] The memory 132 may also include a program / utility 1325 having a set (at least one) of program modules 1324, such program modules 1324 including but not limited to: an operating system, one or more application programs, other program modules, and program data, each of which or some combination may include an implementation of a network environment.

[0214] The electronic device 130 may also communicate with one or more external devices 134 (e.g., keyboards, pointing devices, etc.), may also communicate with one or more devices that enable a user to interact with the electronic device 130, and / or communicate with any device that enables the electronic device 130 to communicate with one or more other electronic devices (e.g., routers, modems, etc.). Such communication may be performed via an input / output (I / O) interface 135. Furthermore, the electronic device 130 may also communicate with one or more networks (e.g., a local area network (LAN), a wide area network (WAN), and / or a public network, such as the Internet) via a network adapter 136. As shown, the network adapter 136 communicates with other modules for the electronic device 130 via a bus 133. It should be understood that, although not shown in the figure, other hardware and / or software modules may be used in conjunction with the electronic device 130, including but not limited to: microcode, device drivers, redundant processors, external disk drive arrays, RAID systems, tape drives, and data backup storage systems.

[0215] In an exemplary embodiment, a computer-readable storage medium including instructions is also provided, such as a memory 132 including instructions, and the instructions can be executed by the processor 131 of the device 400 to complete the above method. Alternatively, the computer-readable storage medium can be a ROM, a random access memory (RAM), a CD-ROM, a magnetic tape, a floppy disk, an optical data storage device, etc.

[0216] In an exemplary embodiment, a computer program product is also provided, including a computer program / instruction, which, when executed by the processor 131, implements any of the methods for designing the Earth-Mars transfer trajectory provided in the present application.

[0217] In an exemplary embodiment, various aspects of a method for designing an Earth-Mars transfer orbit provided by the present application may also be implemented in the form of a program product, which includes program code. When the program product is run on a computer device, the program code is used to enable the computer device to execute the steps of a method for designing an Earth-Mars transfer orbit according to various exemplary embodiments of the present application described above in this specification.

[0218] The program product may employ any combination of one or more readable media. The readable medium may be a readable signal medium or a readable storage medium. The readable storage medium may be, for example, but not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, device or device, or any combination thereof. More specific examples of readable storage media (a non-exhaustive list) include: an electrical connection with one or more wires, a portable disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disk read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination thereof.

[0219] The program product for determining the Earth-Mars transfer trajectory of the embodiment of the present application can adopt a portable compact disk read-only memory (CD-ROM) and include program code, and can be run on an electronic device. However, the program product of the present application is not limited thereto. In this document, a readable storage medium can be any tangible medium containing or storing a program, which can be used by or in combination with an instruction execution system, apparatus, or device.

[0220] A readable signal medium may include a data signal propagated in baseband or as part of a carrier wave, wherein readable program code is carried. Such propagated data signals may take a variety of forms, including, but not limited to, electromagnetic signals, optical signals, or any suitable combination of the foregoing. A readable signal medium may also be any readable medium other than a readable storage medium, which may transmit, propagate, or transfer a program for use by or in conjunction with an instruction execution system, apparatus, or device.

[0221] The program code embodied on the readable medium may be transmitted using any appropriate medium, including but not limited to wireless, wireline, optical fiber cable, RF, etc., or any suitable combination of the foregoing.

[0222] The program code for performing the operations of the present application can be written in any combination of one or more programming languages, including object-oriented programming languages ​​such as Java, C++, etc., and conventional procedural programming languages ​​such as "Like" language or similar programming languages. The program code can be executed entirely on the user electronic device, partially on the user device, as a separate software package, partially on the user electronic device and partially on a remote electronic device, or entirely on a remote electronic device or server. In cases involving remote electronic devices, the remote electronic device can be connected to the user electronic device through any type of network, including a local area network (LAN) or a wide area network (WAN), or can be connected to an external electronic device (for example, using an Internet service provider to connect through the Internet).

[0223] It should be noted that, although several units or subunits of the device are mentioned in the above detailed description, this division is merely exemplary and not mandatory. In fact, according to the embodiments of the present application, the features and functions of two or more units described above can be embodied in one unit. Conversely, the features and functions of one unit described above can be further divided into multiple units to be embodied.

[0224] In addition, although the operations of the method of the present application are described in a specific order in the drawings, this does not require or imply that the operations must be performed in this specific order, or that all the operations shown must be performed to achieve the desired results. Additionally or alternatively, some steps may be omitted, multiple steps may be combined into one step, and / or one step may be decomposed into multiple steps.

[0225] Those skilled in the art will appreciate that the embodiments of the present application may be provided as methods, systems, or computer program products. Therefore, the present application may adopt the form of a complete hardware embodiment, a complete software embodiment, or an embodiment in combination with software and hardware. Moreover, the present application may adopt the form of a computer program product implemented in one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) that include computer-usable program code.

[0226] The present application is described with reference to the flowcharts and / or block diagrams of the methods, devices (systems), and computer program products according to the embodiments of the present application. It should be understood that each process and / or box in the flowchart and / or block diagram, as well as the combination of the processes and / or boxes in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable image scaling device to produce a machine, so that the instructions executed by the processor of the computer or other programmable image scaling device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 A process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0227] These computer program instructions may also be stored in a computer readable memory capable of directing a computer or other programmable image scaling device to operate in a specific manner, so that the instructions stored in the computer readable memory produce an article of manufacture including an instruction device, the instruction device being implemented in the process Figure 1 A process or multiple processes and / or boxes Figure 1 A function specified in one or more boxes.

[0228] These computer program instructions may also be loaded onto a computer or other programmable image scaling device so that a series of operational steps are executed on the computer or other programmable device to produce a computer-implemented process, whereby the instructions executed on the computer or other programmable device provide for implementing the process. Figure 1 A process or multiple processes and / or boxes Figure 1 The steps for the functions specified in one or more boxes.

[0229] Although the preferred embodiments of the present application have been described, those skilled in the art may make additional changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications falling within the scope of the present application.

[0230] Obviously, those skilled in the art can make various changes and modifications to the present application without departing from the spirit and scope of the present application. Thus, if these modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is also intended to include these modifications and variations.

Claims

1. A method for designing an Earth-Mars transfer trajectory, characterized in that: The method comprises: Obtain orbital parameters and preset value ranges for each orbital parameter, wherein the orbital parameters include: the time t0 when the Mars probe enters orbit, the time dt from t0 to the perigee EL , the time from t0 to the deep space maneuver dt DSM , the time from t0 to Mars dt EM , the lunar leverage radius r PL and the lunar leverage azimuth Based on the preset number of iterations, each orbital parameter is iteratively processed by the global optimization algorithm to determine the first minimum velocity increment J from the first total velocity increment J1 of the Mars probe after launch obtained in each iteration. 1min ; wherein the J 1min is the minimum value of J1 obtained in each iteration; Determine the J 1min The first assignment results of the corresponding orbital parameters are used to determine the Earth-Mars transfer orbit based on the first assignment results of the orbital parameters; wherein each iteration process is as follows: For each orbital parameter, assign a value to the orbital parameter based on a preset value interval of the orbital parameter; For each orbital parameter after the assignment, the position vector r of the Mars probe after the lunar leverage is determined based on the orbital integration formula and the principle of leveraged flight. 1+ and the velocity vector v 1+ ; According to the r1+ and the v 1+ Determine the first transfer orbit parameters, and determine the total velocity increment J1 of the Mars probe after launch according to the first transfer orbit parameters, wherein the first transfer orbit parameters include the velocity vector v of the Mars probe after deep space maneuvers 1dsm+ and the velocity vector v when arriving at Mars 1m .

2. The method according to claim 1, characterized in that The position vector r of the Mars probe after the lunar leverage is determined based on the orbital integral formula and the leverage flight principle. 1+ and the velocity vector v 1+ ,include: According to the orbit perigee height r PE , orbital inclination i E And the t0 and dt after assignment EL , taking the orbit transfer angle θ from the entry point to the moon as the iterative variable, determine the initial orbit parameters of the Mars probe rendezvous with the moon, the initial orbit parameters at least include the position vector r of the Mars probe at time t0 01 and the velocity vector v 01 ; Based on the orbital integration formula, determining the position vector r1- and velocity vector v1- of the Mars probe entering the lunar influence sphere according to the initial orbital parameters; Based on the formula of the principle of leveraging flight, according to the r 1- 、The v 1- 、 PL and stated Determine the r 1+ and the v 1+ .

3. The method according to claim 1, characterized in that The first transfer orbit parameter is determined according to the following method: According to the assigned t0 and dt EL Determine the time t1 when the Mars probe arrives at the moon; Based on the orbital integration formula, determine the first probe parameter according to the v1+, the r1+ and the t1, the first probe parameter at least including the speed, position and time of the Mars probe entering the Earth's sphere of influence, and the first probe parameter is the parameter of the Mars probe in the geocentric coordinate system; The first detector parameter is converted based on the cone stitching principle to obtain the second detector parameter, where the second detector parameter is the parameter of the Mars probe in the heliocentric coordinate system; Based on the orbital integral formula, according to the second detector parameter and the assigned dt DSM Determine the time t at which the Mars probe arrives at the deep space maneuvering point 1dsm 、Position vector r 1dsm and the velocity vector v 1dsm- , and according to the t 1dsm The dt EM and the assigned values ​​to query the ephemeris table JPL to determine the time t at which the Mars probe arrives at Mars 1m and the Mars rover at t 1m The position vector r at the time 1m ; Based on Lambert's formula, according to the t 1dsm 、 1dsm , the 1m And the r 1m , to determine the velocity vector v after deep space maneuvering 1dsm+ and the Mars arrival velocity vector v 1m , thereby determining the first transfer orbit parameters.

4. The method according to claim 3, characterized in that Determining the total velocity increment J1 of the Mars probe after launch according to the transfer orbit parameters includes: According to the v 1dsm- 、The v 1dsm+ and the v 1m Determine the velocity increment Δv required to perform a deep space maneuver on the Mars rover 1DSM and the velocity increment Δv required for Mars capture 1MOI ; According to the Δv 1DSM and the Δv 1MOI The sum determines the J1.

5. The method according to claim 1, characterized in that The determining of the Earth-Mars transfer orbit based on the first assignment results of the orbital parameters comprises: For each orbital parameter, the first value assignment result of the orbital parameter is used as the initial value of the orbital parameter, and each orbital parameter is iteratively processed by the step-size acceleration algorithm to determine the second minimum velocity increment J from the second total velocity increment J2 of the Mars probe after launch obtained in each iteration. 2min ; wherein the J 2min is the minimum value of J2 obtained in each iteration; Determine the J 2min The corresponding second assignment results of each orbital parameter are used to determine the Earth-Mars transfer orbit based on the second assignment results of each orbital parameter.

6. The method according to claim 5, characterized in that Each iteration process is as follows: For each orbital parameter after the assignment, the position vector and velocity vector of the Mars probe in the Earth-Mars transfer process are integrated using an accurate perturbation model to obtain the perigee parameters of the Mars probe; wherein the second transfer orbital parameters include the velocity vector v of the Mars probe after the deep space maneuver 2dsm+ And the velocity vector v when reaching the perigee 2m ; The near-fire point parameter is corrected by using a differential correction method so that the corrected near-fire point parameter reaches a preset parameter threshold; The second transfer orbit parameters are corrected based on the corrected perigee parameters, and the J2 is determined according to the corrected second transfer orbit parameters.

7. The method according to claim 6, characterized in that The determining J2 according to the corrected second transfer orbit parameter comprises: Determine the velocity increment Δv required for the Mars probe to perform deep space maneuvers according to the corrected second transfer orbit parameters 2DSM and the velocity increment Δv required for fire capture 2MOI ; According to the Δv 2DSM and the Δv 2MOI The sum determines the J2.

8. A device for designing an Earth-Mars transfer trajectory, characterized in that: The device comprises: The parameter acquisition module is configured to execute the acquisition of orbital parameters and the preset value range of each orbital parameter, wherein the orbital parameters include: the time t0 when the Mars probe enters orbit, the time dt from t0 to the perigee, and the time dt when the Mars probe enters orbit. EL , the time from t0 to the deep space maneuver dt DSM , the time from t0 to Mars dt EM , the lunar leverage radius r PL and the lunar leverage azimuth The velocity increment module is configured to perform iterative processing of each orbit parameter through a global optimization algorithm based on a preset number of iterations, so as to determine a first minimum velocity increment J from the first total velocity increment J1 of the Mars probe after launch obtained in each iteration. 1min ; wherein the J 1min is the minimum value of J1 obtained in each iteration; A trajectory generation module is configured to perform the determination of the J 1min The first assignment results of the corresponding orbital parameters are used to determine the Earth-Mars transfer orbit based on the first assignment results of the orbital parameters; wherein each iteration process is as follows: For each orbital parameter, assign a value to the orbital parameter based on a preset value interval of the orbital parameter; For each orbital parameter after the assignment, the position vector r of the Mars probe after the lunar leverage is determined based on the orbital integration formula and the principle of leveraged flight. 1+ and the velocity vector v 1+ ; According to the r1+ and the v 1+ Determine the first transfer orbit parameters, and determine the total velocity increment J1 of the Mars probe after launch according to the first transfer orbit parameters, wherein the first transfer orbit parameters include the velocity vector v of the Mars probe after deep space maneuvers 1dsm+ and the velocity vector v when arriving at Mars 1m .

9. An electronic device, characterized in that: It comprises at least one processor; and a memory communicatively connected to the at least one processor; wherein the memory stores instructions executable by the at least one processor, and the instructions are executed by the at least one processor so that the at least one processor can execute the method as described in any one of claims 1-7.

10. A computer storage medium, characterized in that: The computer storage medium stores a computer program, and the computer program is used to enable a computer to execute the method according to any one of claims 1 to 7.

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