A design method and related device for a probe free return orbit

Abstract

The embodiment of the present application provides a design method and related device for a free return orbit of a detector, by iteratively processing the initial value of each orbit design parameter based on the target parameter required by the mission constraint, so as to determine the iterative deviation of the P-plane parameter of the detector. And based on the iterative deviation, the orbit design parameter is corrected to determine a free return orbit of different orbit configurations that meets the target parameter. In the iterative process, based on the B-plane parameter formula, the first parameter value of the P-plane parameter is determined according to the vector parameter of the detector, and the second parameter value of the P-plane parameter is determined according to the vector parameter and the target parameter, and then the iterative deviation is determined according to the first parameter value and the second parameter value. For each iterative process, if the iterative deviation is greater than the preset threshold, the initial value is corrected based on the differential correction method, and the corrected orbit design parameters are used as the initial value of the next iteration, and the iterative process is repeated until the iterative deviation is no greater than the preset threshold.

Description

Technical Field

[0001] The present invention relates to the field of deep space exploration technology, and in particular to a design method for a detector free return 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 field of aerospace. Deep space exploration can not only enhance human 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. The moon is the closest natural planet to the earth and an ideal transit station to deep space. As one of the key technologies for manned lunar missions, the design of manned lunar orbit has always been a hot research topic in this field. Unlike unmanned exploration, the manned lunar orbit must ensure the safety of astronauts and must have the ability to safely return to the ground when failures occur at various stages of the mission. The free return orbit is an orbit that starts from the near earth and flies to the moon, and then returns to the earth without power after deceleration by the moon. It has high safety.

[0003] In the related art, there are 16 orbital configurations for free return orbits. The traditional orbital design scheme conducts preliminary design under a simplified model for one or several of these configurations, and then performs a revised design under a high-precision model. This design method requires layer-by-layer revision of orbital design parameters, has low computational efficiency and is not universal. In addition, the above design method has the problem of incomplete consideration of the task constraints of engineering implementation, and has certain limitations. Summary of the invention

[0004] An embodiment of the present application provides a design method and related device for a free return orbit of a probe, which determines the iterative deviation of the P-plane parameters of the probe based on target parameters required by mission constraints, and corrects the orbit design parameters based on the iterative deviation to determine a free return orbit of a variety of orbital configurations that meet the target parameters.

[0005] In a first aspect, an embodiment of the present application provides a method for designing a free return trajectory of a probe, the method comprising:

[0006] The initial values ​​of each orbital design parameter are iteratively processed based on the target parameter, so as to correct each orbital design parameter according to the iterative deviation obtained in each iteration, and a free return orbit is constructed according to each orbital parameter that meets the engineering requirements after correction; wherein the orbital design parameter at least includes the position component z of the probe on the z-axis of the preset inertial system 0 and the velocity components v on the x, y, and z axes of the preset inertial system x 、v y 、v zThe preset inertial system is a pre-constructed ecliptic plane symmetrical free return orbit instantaneous ecliptic coordinate system of the moon center at the perigee time; the target parameters include at least the first target parameter or the second target parameter; the first target parameter includes the orbital inclination i of the entry point and the target perigee distance r p The second target parameters include the reentry angle γ and the reentry point geocentric distance r e ;

[0007] The iterative processing process is as follows:

[0008] In the first iteration, the initial value is determined based on the pre-constructed symmetric free return orbit in the ecliptic plane, and the vector parameters corresponding to the initial value are determined by dynamic integration; wherein the vector parameters at least include the velocity vector v and the position vector r of the probe at the perigee or the velocity vector v and the position vector r of the probe at the re-entry point;

[0009] Based on the B-plane parameter formula, a first parameter value of the P-plane parameter of the detector is determined according to the vector parameter, and a second parameter value of the P-plane parameter is determined according to the target parameter and the vector parameter; wherein, if the orbit to be determined is a hyperbolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the orbit asymptote, and if it is an elliptical orbit or a parabolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the pericenter vector;

[0010] Determine an iteration deviation of this iteration according to the first parameter value and the second parameter value;

[0011] For each iterative processing, if the iterative deviation is greater than a preset threshold, the initial value is corrected based on the differential correction method to obtain the corrected value of each track design parameter, and the corrected value is used as the initial value of each track design parameter in the next iteration, and the iterative process is repeated until the iterative deviation is no greater than the preset threshold.

[0012] The embodiment of the present application iterates the initial value of each orbital design parameter by using the target parameter based on the task constraint requirement, so as to correct each orbital design parameter according to the iterative deviation obtained in each iteration, and construct a free return orbit according to each orbital parameter that meets the engineering requirements after correction. Compared with the traditional B-plane parameters, the P-plane parameters are applicable to more orbital configurations. In the above-mentioned iterative process, it is necessary to determine the first parameter value of the P-plane parameter of the detector based on the B-plane parameter formula through the vector parameter of the detector. And determine the second parameter value of the P-plane parameter based on the target parameter and the vector parameter. Then determine the iterative deviation of this iteration based on the first parameter value and the second parameter value. For each iterative process, if the iterative deviation is greater than the preset threshold, the initial value is corrected based on the differential correction method to obtain the corrected value of each orbital design parameter, and the corrected value is used as the initial value of each orbital design parameter in the next iteration, and the iterative process is repeated until the iterative deviation is no greater than the preset threshold.

[0013] The above process determines the iterative deviation of the P-plane parameters of the probe based on the target parameters of the mission constraint requirements, and corrects the orbit design parameters based on the iterative deviation to determine the free return orbits of various orbital configurations that meet the target parameters.

[0014] In some possible embodiments, the symmetrical free return orbit in the ecliptic plane is constructed with the perigee of the probe as the starting point; the starting point is located on the x-axis of the preset inertial system, and the preset inertial system is parallel to the instantaneous ecliptic coordinate system of the moon center at the perigee moment of the symmetrical free return orbit in the ecliptic plane; the earth-moon transfer orbit and the moon-earth return orbit of the symmetrical free return orbit in the ecliptic plane are symmetrical on both sides of the x-axis;

[0015] In the first iteration, the initial values ​​of the orbit design parameters are determined based on the pre-constructed symmetrical free return orbit in the orbit plane, including:

[0016] Determine z according to the orbit structure of the symmetrical free return orbit in the ecliptic plane 0 、v x and v z The initial value of z 0 、v x and v z The value in the symmetric free return orbit model in the ecliptic plane is 0;

[0017] Based on the preset speed interval, the detector performs multiple acquisition operations on the parameters to be processed in the symmetrical free return orbit in the ecliptic plane; wherein the parameters to be processed at least include the position component x of the detector on the x-axis at each acquisition. 0 and the velocity component v on the y-axis 0 ;

[0018] After each acquisition, the pericentric distance r corresponding to the current acquisition operation is determined based on the parameters to be processed. pe , and according to the r corresponding to each acquisition operation pe Determine the v y The initial value of .

[0019] The symmetric free return orbit in the ecliptic plane of the present application embodiment is constructed with the perigee of the probe as the starting point; the starting point is located at the x-axis of the preset inertial system, and the preset inertial system is parallel to the instantaneous ecliptic coordinate system of the moon center at the perigee of the symmetric free return orbit in the ecliptic plane; the earth-moon transfer orbit and the moon-earth return orbit of the symmetric free return orbit in the ecliptic plane are symmetrical on both sides of the x-axis. In the symmetric free return orbit in the ecliptic plane obtained by the above construction method, z 0 、v x and v z The value of is always 0. Therefore, when determining the initial values ​​of each orbit design parameter in the first iteration, z can be determined based on the orbit structure of the symmetrical free return orbit in the ecliptic plane. 0 、v x and v z Determine the initial value of v y When the initial value is set, multiple acquisition operations are performed on the parameters to be processed in the symmetrical free return orbit of the detector in the ecliptic plane based on the preset speed interval; after each acquisition, the pericentric distance r corresponding to this round of acquisition operation is determined based on the parameters to be processed. pe , and according to the r corresponding to each acquisition operation pe Determine v y The initial value of .

[0020] In some possible embodiments, the determining of the pericentric distance r corresponding to the current round of acquisition operation based on the parameters to be processed is pe ,include:

[0021] According to the x 0 and the v 0 Determine the semi-major axis a, the eccentricity e and the semi-diameter p of the orbit, and determine the true anomaly θ of the boundary point of the lunar influence sphere based on a, e and p;

[0022] Determine the radial velocity component v of the probe at the boundary of the lunar sphere of influence according to θ r , vertical radial velocity component v u and the flight time Δt, and according to the v r 、The v u and the Δt to determine the position vector R of the moon in the preset inertial system L and velocity vector V L ;

[0023] According to the RL and the V L Determine the position vector R of the detector at the center of the earth e and velocity vector V e , and according to the R e and the V e Determine the pericentric distance r corresponding to this round of acquisition operation pe .

[0024] The present application embodiment is based on x 0 and v 0 Determine the semi-major axis a, eccentricity e, and orbital semi-diameter p, and determine the true anomaly θ at the boundary of the lunar influence sphere based on a, e, and p. Then determine the radial velocity component v of the probe at the boundary of the lunar influence sphere based on θ. r , vertical radial velocity component v u and the flight time Δt, and according to v r 、v u and Δt determine the position vector R of the moon in the preset inertial system L and velocity vector V L Finally, according to R L and V L Determine the position vector R of the detector at the center of the earth e and velocity vector V e , according to R e and V e Determine the pericentric distance r corresponding to this round of acquisition operation pe .

[0025] In some possible embodiments, determining the first parameter value of the P-plane parameter of the detector according to the vector parameter based on the B-plane parameter formula includes:

[0026] Based on the B-plane parameter formula, the n vector, the orbit semi-radius p and the E vector are determined according to the vector parameters; wherein the n vector is the orbit normal vector; and the E vector is the pericenter vector;

[0027] Based on the orbit type of the free return orbit, according to the n vector, the p 2 And the E vector determines the P vector, and the first parameter value is determined according to the P vector.

[0028] The P plane parameters in the embodiment of the present application can be applied to free return orbits of different configurations. When determining the first parameter value, based on the orbit type of the free return orbit, according to the n vector, p 2 And the E vector determines the P vector, and the first parameter value is determined according to the P vector.

[0029] In some possible embodiments, the n vector, the p2 And the E vector determines the P vector, including:

[0030] If the orbit type is a hyperbolic orbit type, the S vector is determined according to the n vector and the E vector, and the P vector is determined according to the p, the S vector and the n vector; wherein, for the hyperbolic orbit type, the S vector is a unit vector passing through the center of the reference celestial body and having a direction parallel to the asymptote of the hyperbolic orbit, and the vector direction of the P vector is from the center of the reference celestial body to the intersection of the asymptote of the orbit and the P plane;

[0031] If the orbit type is the elliptical orbit type or the parabolic orbit type, the P vector is determined according to the p, the E vector and the n vector; wherein, for the elliptical orbit type or the parabolic orbit type, the vector direction of the P vector is perpendicular to the E vector from the center of the reference celestial body and points to the side entering the orbit.

[0032] The P plane parameters of the embodiment of the present application can be applied to different orbit configurations. For a hyperbolic orbit, the S vector can be determined according to the n vector and the E vector, and the P vector can be determined according to the p, S vector and the n vector. For an elliptical orbit or a parabolic orbit, the P vector can be determined according to the p, E vector and the n vector.

[0033] In some possible embodiments, determining the first parameter value according to the P vector includes:

[0034] If the track type is a hyperbolic track type, a T vector and an R vector are determined according to the S vector and the N vector; wherein the T vector and the R vector are unit vectors perpendicular to the P plane;

[0035] If the orbit type is an elliptical orbit type or a parabolic orbit type, the T vector and the R vector are determined according to the E vector and the N vector; wherein the N vector is a reference vector determined according to the rotation axis of a reference celestial body.

[0036] The first parameter value is determined according to the T vector, the R vector, and the P vector.

[0037] The P-plane parameters of the embodiment of the present application can be applied to different orbit configurations. For a hyperbolic orbit, the T-vector and the R-vector can be determined according to the S-vector and the N-vector. For an elliptical orbit or a parabolic orbit, the T-vector and the R-vector can be determined according to the E-vector and the N-vector. Finally, the first parameter value is determined according to the T-vector, the R-vector and the P-vector. The first parameter value can characterize the P-plane parameter value of the detector without considering the target parameters.

[0038] In some possible embodiments, determining the second parameter value of the P-plane parameter according to the target parameter and the vector parameter includes:

[0039] Determine the semi-major axis a according to the vector parameter;

[0040] Determine the angle α based on the orbit type of the free return orbit, and determine the orbit semi-path p according to the target parameter and a; wherein the angle α is the angle between the P plane and the reference plane; the reference plane is a plane passing through the center of the reference celestial body and perpendicular to the n vector;

[0041] Based on the right spherical triangle formula, the angle is determined according to the target parameter and α Among them, the angle is the angle between the P vector and the T vector;

[0042] According to the p and the angle The second parameter value is determined.

[0043] In the embodiment of the present application, when determining the second parameter value, the semi-major axis a is determined according to the vector parameter, and then the angle α is determined based on the orbit type of the free return orbit, and the orbit semi-path p is determined according to the target parameter and a. Finally, the angle is determined according to the right spherical triangle formula according to the target parameter and α. And according to p and the angle Determine a second parameter value. The second parameter value can represent the P-plane parameter value of the detector in combination with the target parameter, and the target parameter is the task constraint in the engineering task, so the free return orbit of the present application can be applied to the target parameters of different task constraints.

[0044] In some possible embodiments, determining the angle α based on the orbit type of the free return orbit includes:

[0045] If the orbit type is the hyperbola type, determining the S vector according to the n vector and the E vector, and determining the angle α according to the component of the S vector in the z-axis direction of the preset inertial system;

[0046] If the orbit type is the ellipse type or the parabola type, the angle α is determined according to the component of the E vector in the z-axis direction of the preset inertial system.

[0047] The P plane parameters of the embodiment of the present application can be applied to different orbit configurations. When determining the angle α, for a hyperbolic orbit, the S vector can be determined according to the n vector and the E vector, and the angle α can be determined according to the component of the S vector in the z-axis direction of the preset inertial system. For an elliptical orbit or a parabolic orbit, the S vector can be determined according to the n vector and the E vector, and the angle α can be determined according to the component of the S vector in the z-axis direction of the preset inertial system.

[0048] In a second aspect, an embodiment of the present application provides a device for designing a free return trajectory of a probe, the device comprising:

[0049] A parameter determination module is configured to perform iterative processing on the initial values ​​of each track design parameter based on the target parameter, so as to correct each track design parameter according to an iterative deviation obtained in each iteration;

[0050] The orbit determination module is configured to construct a free return orbit according to the orbit parameters that meet the engineering requirements after correction; wherein the orbit design parameters at least include the position component z of the probe on the z-axis of the preset inertial system 0 and the velocity components v on the x, y, and z axes of the preset inertial system x 、v y 、v z The preset inertial system is a pre-constructed ecliptic plane symmetrical free return orbit instantaneous ecliptic coordinate system of the moon center at the perigee time; the target parameters include at least the first target parameter or the second target parameter; the first target parameter includes the orbital inclination i of the entry point and the target perigee distance r p The second target parameters include the reentry angle γ and the reentry point geocentric distance r e ;

[0051] The iterative processing process is as follows:

[0052] In the first iteration, the initial value is determined based on the pre-constructed symmetric free return orbit in the ecliptic plane, and the vector parameters corresponding to the initial value are determined by dynamic integration; wherein the vector parameters at least include the velocity vector v and the position vector r of the probe at the perigee or the velocity vector v and the position vector r of the probe at the re-entry point;

[0053] Based on the B-plane parameter formula, a first parameter value of the P-plane parameter of the detector is determined according to the vector parameter, and a second parameter value of the P-plane parameter is determined according to the target parameter and the vector parameter; wherein, if the orbit to be determined is a hyperbolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the orbit asymptote, and if it is an elliptical orbit or a parabolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the pericenter vector;

[0054] Determine an iteration deviation of this iteration according to the first parameter value and the second parameter value;

[0055] For each iterative processing, if the iterative deviation is greater than a preset threshold, the initial value is corrected based on the differential correction method to obtain the corrected value of each track design parameter, and the corrected value is used as the initial value of each track design parameter in the next iteration, and the iterative process is repeated until the iterative deviation is no greater than the preset threshold.

[0056] In some possible embodiments, the symmetrical free return orbit in the ecliptic plane is constructed with the perigee of the probe as the starting point; the starting point is located on the x-axis of the preset inertial system, and the preset inertial system is parallel to the instantaneous ecliptic coordinate system of the moon center at the perigee moment of the symmetrical free return orbit in the ecliptic plane; the earth-moon transfer orbit and the moon-earth return orbit of the symmetrical free return orbit in the ecliptic plane are symmetrical on both sides of the x-axis;

[0057] When performing the first iteration, the initial values ​​of the track design parameters are determined based on the pre-constructed symmetrical free return track in the white track plane, and the parameter determination module is configured as follows:

[0058] Determine z according to the orbit structure of the symmetrical free return orbit in the ecliptic plane 0 、v x and v z The initial value of z 0 、v x and v z The value in the symmetric free return orbit model in the ecliptic plane is 0;

[0059] Based on the preset speed interval, the detector performs multiple acquisition operations on the parameters to be processed in the symmetrical free return orbit in the ecliptic plane; wherein the parameters to be processed at least include the position component x of the detector on the x-axis at each acquisition. 0 and the velocity component v on the y-axis 0 ;

[0060] After each acquisition, the pericentric distance r corresponding to the current acquisition operation is determined based on the parameters to be processed. pe , and according to the r corresponding to each acquisition operation pe Determine the v y The initial value of .

[0061] In some possible embodiments, the step of determining the pericentric distance r corresponding to the current round of acquisition operation based on the parameters to be processed is performed. pe , the parameter determination module is configured as follows:

[0062] According to the x 0 and the v 0 Determine the semi-major axis a, the eccentricity e and the semi-diameter p of the orbit, and determine the true anomaly θ of the boundary point of the lunar influence sphere based on a, e and p;

[0063] Determine the radial velocity component v of the probe at the boundary of the lunar sphere of influence according to θ r , vertical radial velocity component v u and the flight time Δt, and according to the vr 、The v u and the Δt to determine the position vector R of the moon in the preset inertial system L and velocity vector V L ;

[0064] According to the R L and the V L Determine the position vector R of the detector at the center of the earth e and velocity vector V e , and according to the R e and the V e Determine the pericentric distance r corresponding to this round of acquisition operation pe .

[0065] In some possible embodiments, the B-plane parameter formula is executed to determine the first parameter value of the P-plane parameter of the detector according to the vector parameter, and the parameter determination module is configured as follows:

[0066] Based on the B-plane parameter formula, the n vector, the orbit semi-radius p and the E vector are determined according to the vector parameters; wherein the n vector is the orbit normal vector; and the E vector is the pericenter vector;

[0067] Based on the orbit type of the free return orbit, according to the n vector, the p 2 And the E vector determines the P vector, and the first parameter value is determined according to the P vector.

[0068] In some possible embodiments, performing the step of: 2 And the E vector determines the P vector, and the parameter determination module is configured as follows:

[0069] If the orbit type is a hyperbolic orbit type, the S vector is determined according to the n vector and the E vector, and the P vector is determined according to the p, the S vector and the n vector; wherein, for the hyperbolic orbit type, the S vector is a unit vector passing through the center of the reference celestial body and having a direction parallel to the asymptote of the hyperbolic orbit, and the vector direction of the P vector is from the center of the reference celestial body to the intersection of the asymptote of the orbit and the P plane;

[0070] If the orbit type is the elliptical orbit type or the parabolic orbit type, the P vector is determined according to the p, the E vector and the n vector; wherein, for the elliptical orbit type or the parabolic orbit type, the vector direction of the P vector is perpendicular to the E vector from the center of the reference celestial body and points to the side entering the orbit.

[0071] In some possible embodiments, the determining of the first parameter value according to the P vector is performed, and the parameter determination module is configured to:

[0072] If the track type is a hyperbolic track type, a T vector and an R vector are determined according to the S vector and the N vector; wherein the T vector and the R vector are unit vectors perpendicular to the P plane;

[0073] If the orbit type is an elliptical orbit type or a parabolic orbit type, the T vector and the R vector are determined according to the E vector and the N vector; wherein the N vector is a reference vector determined according to the rotation axis of a reference celestial body.

[0074] The first parameter value is determined according to the T vector, the R vector, and the P vector.

[0075] In some possible embodiments, the determining of the second parameter value of the P-plane parameter according to the target parameter and the vector parameter is performed, and the parameter determination module is configured as follows:

[0076] Determine the semi-major axis a according to the vector parameter;

[0077] Determine the angle α based on the orbit type of the free return orbit, and determine the orbit semi-path p according to the target parameter and a; wherein the angle α is the angle between the P plane and the reference plane; the reference plane is a plane passing through the center of the reference celestial body and perpendicular to the n vector;

[0078] Based on the right spherical triangle formula, the angle is determined according to the target parameter and α Among them, the angle is the angle between the P vector and the T vector;

[0079] According to the p and the angle The second parameter value is determined.

[0080] In some possible embodiments, the determining of the angle α based on the track type of the free return track is performed, and the parameter determination module is configured as follows:

[0081] If the orbit type is the hyperbola type, determining the S vector according to the n vector and the E vector, and determining the angle α according to the component of the S vector in the z-axis direction of the preset inertial system;

[0082] If the orbit type is the ellipse type or the parabola type, the angle α is determined according to the component of the E vector in the z-axis direction of the preset inertial system.

[0083] In a third aspect, an embodiment of the present application further provides an electronic device, including:

[0084] processor;

[0085] a memory for storing instructions executable by the processor;

[0086] The processor is configured to execute the instructions to implement any one of the methods provided in the first aspect of the present application.

[0087] 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.

[0088] 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

[0089] 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.

[0090] Figure 1a This is an overall flow chart of the design method of the detector free return trajectory shown in the embodiment of the present application;

[0091] Figure 1b A schematic diagram of B-plane parameters shown in an embodiment of the present application;

[0092] Figure 1c A schematic diagram of P plane parameters of a hyperbolic track type shown in an embodiment of the present application;

[0093] Figure 1d A schematic diagram of P plane parameters of an elliptical or parabolic orbit type shown in an embodiment of the present application;

[0094] Figure 2a A flowchart of an iterative process shown in an embodiment of the present application;

[0095] Figure 2b This is a schematic diagram of a symmetrical free return track within the orbital plane shown in an embodiment of the present application;

[0096] Figure 2cThis is a schematic diagram of a symmetrical free return orbit state within the orbital plane shown in an embodiment of the present application;

[0097] Figure 2d Different x shown in the embodiment of this application 0 The corresponding r pe With v y Schematic diagram of the relationship between

[0098] Figure 2e A schematic diagram of solving the second parameter value of the P plane parameter shown in an embodiment of the present application;

[0099] Figure 3a A schematic diagram of free return orbit parameters under a double two-body model shown in an embodiment of the present application;

[0100] Figure 3b A schematic diagram of free return orbit parameters under a high-precision model shown in an embodiment of the present application;

[0101] Figure 3c The three-dimensional views of the free return orbits of the four configurations of the lunar retrograde motion shown in the embodiments of the present application;

[0102] Figure 3d The three-dimensional views of the free return orbits of the four configurations of the lunar center prograde shown in the embodiments of the present application;

[0103] Figure 3e The transfer time and |x shown in the embodiment of the present application 0 | Schematic diagram of relationship curve;

[0104] Figure 4 A block diagram of a design device 400 for a probe free return trajectory according to an embodiment of the present application;

[0105] Figure 5 This is a schematic diagram of an electronic device shown in an embodiment of the present application. DETAILED DESCRIPTION

[0106] 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, " / " means 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.

[0107] 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.

[0108] 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.

[0109] In the related technology, the probe is launched from the ground by a carrier rocket to a low-altitude Earth parking orbit. After the parking orbit glides to a suitable position, the engine starts to accelerate the probe into the Earth-Moon transfer orbit. During the flight of the probe, orbital errors are corrected, and when it reaches the vicinity of the moon, the braking engine starts to work. At this time, the probe slows down and is captured by the moon, becoming a lunar satellite. After completing its circumlunar mission, the probe needs to slow down at a suitable position to enter the lunar landing orbit. Traditional orbital design schemes are mostly aimed at one or several known orbital configurations (there are 16 orbital configurations in the free return orbit), and the B-plane parameter algorithm is used to preliminarily design the orbital design parameters under a simplified model, and then the design is corrected in a high-precision model. The above design method requires layer-by-layer correction of the orbital design parameters, which has low computational efficiency and is not universal.

[0110] To solve the above problems, the inventive concept of the present application is: the initial values ​​of each orbital design parameter are iteratively processed by the target parameters based on the task constraint requirements, so as to correct each orbital design parameter according to the iterative deviation obtained in each iteration, and construct a free return orbit according to each orbital parameter that meets the engineering requirements after correction. Compared with the traditional B-plane parameters, the P-plane parameters are applicable to more orbital configurations. In the above iteration process, it is necessary to determine the first parameter value of the P-plane parameter of the detector based on the B-plane parameter formula through the vector parameter of the detector. And determine the second parameter value of the P-plane parameter based on the target parameter and the vector parameter. Then determine the iteration deviation of this iteration based on the first parameter value and the second parameter value. For each iteration process, if the iteration deviation is greater than the preset threshold, the initial value is corrected based on the differential correction method to obtain the corrected value of each orbital design parameter, and the corrected value is used as the initial value of each orbital design parameter in the next iteration, and the iteration process is repeated until the iteration deviation is no greater than the preset threshold.

[0111] The above process determines the iterative deviation of the P-plane parameters of the probe based on the target parameters of the mission constraint requirements, and corrects the orbit design parameters based on the iterative deviation to determine the free return orbits of various orbital configurations that meet the target parameters.

[0112] To facilitate understanding of the technical solution provided in the embodiment of the present application, a design method for a detector free return track provided in the embodiment of the present application is described in detail below in conjunction with the accompanying drawings. Figure 1a As shown, the following steps are included:

[0113] Step 101: iteratively processing the initial values ​​of each track design parameter based on the target parameter, so as to correct each track design parameter according to the iterative deviation obtained in each iteration;

[0114] Step 102: construct a free return orbit according to the corrected orbit parameters that meet the engineering requirements; wherein the orbit design parameters at least include the position component z of the probe on the z-axis of the preset inertial system 0 and the velocity components v on the x, y, and z axes of the preset inertial system x 、v y 、v z The preset inertial system is a pre-constructed ecliptic plane symmetrical free return orbit instantaneous ecliptic coordinate system of the moon center at the perigee time; the target parameters include at least the first target parameter or the second target parameter; the first target parameter includes the orbital inclination i of the entry point and the target perigee distance r p The second target parameters include the reentry angle γ and the reentry point geocentric distance r e .

[0115] As mentioned above, traditional orbit design schemes mostly use the B-plane parameter algorithm to make a preliminary design of orbit design parameters under a simplified model, and then make a correction design under a high-precision model. Usually, the speed of a probe launched from the Earth to the vicinity of the Moon exceeds the escape velocity of the probe relative to the Moon at that location. Therefore, the probe will fly in a hyperbolic orbit relative to the center of the Moon near the Moon. Figure 1b As shown, it is a plane passing through the center of the moon and perpendicular to the hyperbola's infinite velocity. Figure 1b Where S is a unit vector passing through the center of the moon and parallel to the asymptote of the entry orbit; T and R are unit vectors that are perpendicular to each other in the B plane; N is the normal unit vector of the lunar orbital plane; B is a vector pointing from the center of the moon to the intersection of the asymptote of the entry orbit and the B plane. Related technologies are mostly based on the B-plane parameter formula, and the parameter values ​​of the B-plane parameters are determined by the above parameters, and the orbit design parameters are corrected according to the B-plane parameters to determine the free return orbit of the probe. Among them, the B-plane parameters are Figure 1bThe projection of the B vector under R and T is shown in .

[0116] In addition to the hyperbolic orbit type, the existing orbit types also include elliptical orbit types, parabolic orbit types, and many others. The B-plane parameters are not suitable for the design of elliptical orbit types. In addition, although some people have proposed the concept of an elliptical B-plane, which can also be adapted to elliptical orbits, the elliptical B-plane parameters are discontinuous. For orbits with relatively large eccentricities, such as the lunar-earth transfer orbit, the eccentricity may exceed 1 during the solution, which may easily lead to non-convergence. Based on this, the present application constructs the P-plane in a similar manner to the B-plane. Compared with the B-plane parameters, the P-plane parameters can be applied to elliptical orbits, and the above-mentioned elliptical B-plane parameters do not exist.

[0117] The embodiment of the present application determines the P plane parameters of the detector in the P plane based on the B plane parameter formula. Specifically, Figure 1c The P plane parameter definition method for the hyperbolic track type is shown. Figure 1c As shown, the P plane is a plane passing through the center of the reference celestial body and perpendicular to the orbit asymptote. The S vector is a unit vector passing through the center of the reference celestial body and parallel to the asymptote of the hyperbolic orbit. Figure 1d The P-plane parameter definition method for elliptical or hyperbolic orbit types is shown. Figure 1d As shown in the figure, the P plane is a plane passing through the center of the reference celestial body and perpendicular to the pericenter. The E vector is the pericenter vector. Figure 1c and Figure 1d The P plane parameters in represent the projection of the P vector on the R vector and the T vector. The R vector and the N vector are unit vectors perpendicular to the P plane. If the free return orbit to be determined is a hyperbolic orbit, the vector direction of the P vector is from the center of the reference celestial body to the intersection of the asymptote of the entry orbit and the P plane; if the free return orbit to be determined is an elliptical orbit type or a hyperbolic orbit type, the vector direction of the P vector is perpendicular to the E vector from the center of the reference celestial body and points to the side of the entry orbit.

[0118] The definition of the above P-plane parameters is similar to that of the B-plane parameters, except that the modulus of the B-vector is replaced by the orbital semi-diameter p instead of the minor semi-axis b. For elliptical and parabolic orbits that are not applicable to the B-plane, Figure 1d The E vector shown in Figure 1c Thus, the parameter values ​​of the above-mentioned P-plane parameters can be determined based on the B-plane parameter formula.

[0119] To facilitate understanding of the above process, Figure 2a As shown, Figure 2a The iterative process in the above step 101 is shown, comprising the following steps:

[0120] Step 201: During the first iteration, the initial value is determined based on the pre-constructed symmetric free return orbit in the ecliptic plane, and the vector parameters corresponding to the initial value are determined by dynamic integration; wherein the vector parameters at least include the velocity vector v and the position vector r of the probe at the perigee or the velocity vector v and the position vector r of the probe at the re-entry point;

[0121] The symmetrical free return orbit in the ecliptic plane is the simplest type of free return orbit. It can be used as a good initial value when solving the three-dimensional free return orbit, and can also be used for preliminary orbit analysis. Figure 2b As shown, the symmetrical free return orbit in the ecliptic plane of the embodiment of the present application is constructed with the perigee of the probe as the starting point, as shown in FIG. Figure 2b As shown in the figure, point A indicates the position where the probe enters the lunar influence sphere from the earth; point B indicates the position where the probe exits the lunar influence sphere from the moon. The starting point C is located on the x-axis of the preset inertial system, which is parallel to the instantaneous ecliptic coordinate system of the moon center at the perigee of the symmetrical free return orbit in the ecliptic plane. The earth-moon transfer orbit and the moon-earth return orbit of the symmetrical free return orbit in the ecliptic plane are symmetrical on both sides of the x-axis.

[0122] When executing the above step 201, z can be determined according to the orbit structure of the symmetrical free return orbit in the ecliptic plane. 0 、v x and v z The initial value of . Specifically, Figure 2b The plane where the xy axis is shown is the ecliptic plane (i.e. there is no z axis). Therefore, the parameter values ​​of the track outside the plane are all 0, so the z in the above track design parameters is 0 and v z The value of is always 0. In addition, since the orbit solved here is symmetric about the x-axis, the initial point C is located on the symmetry axis, so v x The value of is also 0. 0 、v x and v z After the initial value of y The initial value of .

[0123] During implementation, multiple acquisition operations are performed on the parameters to be processed of the detector in the symmetrical free return orbit in the ecliptic plane based on the preset speed interval; wherein the parameters to be processed at least include the position component x of the detector on the x-axis at each acquisition. 0 and the velocity component v on the y-axis 0 ; After each acquisition, the pericentric distance r corresponding to the current acquisition operation is determined based on the parameters to be processed pe , and according to the r corresponding to each acquisition operation pe Determine v y The initial value of .

[0124] Specific examples include Figure 2c As shown, first, according to the position component x of the starting point C on the x-axis 0 and the velocity component v in the y-axis direction 0 , the semi-major axis a, eccentricity e and orbital semi-diameter p are determined by the following formula (1):

[0125]

[0126] Where a is the semi-major axis, e is the eccentricity, p is the semi-diameter of the track, μ L is the lunar gravitational constant. Further, the true anomaly θ of the lunar influence sphere boundary point is determined according to a, e and p by the following formula (2):

[0127]

[0128] Among them, R s is the radius of the lunar sphere of influence; θ is the true anomaly angle, and a positive value represents the state of the exit point. Further, according to θ, the radial velocity component v of the probe at the boundary of the lunar sphere of influence can be determined: r , vertical radial velocity component v u and the flight time Δt; as shown in the following formulas (3) and (4):

[0129]

[0130] Among them, v r is the radial velocity component of the probe at the boundary of the lunar sphere of influence, v u is the vertical radial velocity component of the probe at the boundary of the lunar influence sphere.

[0131]

[0132] Among them, E A is the anomaly angle, Δt is the flight time. r 、v u and Δt can be used to determine the position vector R of the moon in the preset inertial system through the above formulas (5) and (6): L and velocity vector V L :

[0133] β=ω L Δt formula (5)

[0134] Among them, β is the angle of the moon's motion, ω L is the average angular velocity of the moon's revolution.

[0135]

[0136] Among them, L is the distance between the earth and the moon, and V is the speed of the moon orbiting the earth. L and V L , the position vector R of the detector at the center of the earth is determined by the following formulas (7) to (9): e and velocity vector V e :

[0137] When the starting point C is outside the moon, the following formula (7) is used to determine the position vector R of the probe in the preset inertial system at the boundary of the influence sphere: m and velocity vector V m :

[0138]

[0139] When the starting point C is on the inner side of the moon, the following formula (8) is used to determine the position vector R of the probe in the preset inertial system at the boundary of the influence sphere: m and velocity vector V m :

[0140]

[0141] Determine the position vector R of the detector at the center of the earth according to the following formula (9): e and velocity vector V e :

[0142]

[0143] According to R e and V e , the pericentric distance r corresponding to the current round of acquisition operation is determined by the following formulas (10) to (13): pe :

[0144]

[0145] Among them, V eu is the lateral velocity component relative to the center of the earth; V er is the radial velocity component relative to the center of the earth; V ex and V ey R e Components in the x and y directions.

[0146] p e =V eu 2 |R e | 2 / μ Formula (11)

[0147] Among them, p e is the orbital semi-diameter, and μ is the gravitational constant of the Earth.

[0148]

[0149] Among them, E s is esinθ; E c is ecosθ; μ e is the gravitational constant.

[0150]

[0151] The above formulas (1) to (13) can be used to determine the near-center distance r corresponding to each acquisition operation: pe , further, according to the r corresponding to each acquisition operation pe Determine v y It should be understood that the analysis method for entering the influence sphere boundary is the same as that for exiting the influence sphere. For ease of explanation, the case where the detector exits the influence sphere is used as an example. Figure 2d As shown, Figure 2d When the detector produces an impact ball, it is different x 0 The corresponding r pe With v y The relationship between them.

[0152] Figure 2d The starting point of each curve in v y The distance from the center of the moon is x 0 The lunar escape velocity at . Figure 2d It can be seen that the minimum value of each curve is 0, so for a given starting point C, the x 0 Position and target perigee height r pe There are usually two solutions, one with a larger speed and the other with a smaller speed. For a free return orbit, when the probe reaches the perigee, the angle between the speed relative to the moon and the speed of the moon's revolution is usually greater than 90 degrees, so v y It is always negative. Therefore, the solution with a large speed has a speed at the center of the earth opposite to the direction of the moon's movement after leaving the impact sphere, and is in a retrograde orbit relative to the center of the earth.

[0153] In addition, the curve in the figure is a single valley function. The quadratic difference method or 618 method and other single variable optimization algorithms can quickly solve the v with a perigee height of 0. 0 , the left boundary is the lunar escape velocity, and the right boundary can be set to 5000. Then use v 0 As the initial value, the direct and retrograde solutions can be obtained by using nonlinear equation solving methods such as Brent's method or Newton's method. y Request r pe The calculation process is very simple and the amount of calculation is very small. It takes only about 1 microsecond to calculate once. To solve a free return orbit in the ecliptic plane, it is usually only necessary to calculate v yThe calculation of rpe is about 30 times, which takes about 0.03 milliseconds, which is almost negligible. 0 When x is positive, the orbit relative to the center of the moon is a retrograde orbit. 0 When it is negative, the orbit relative to the center of the moon is a prograde orbit. Therefore, there are four configurations of free return orbits in the ecliptic plane, namely: geocentric prograde and mooncentric prograde, geocentric prograde and mooncentric retrograde, geocentric retrograde and mooncentric prograde, and geocentric retrograde and mooncentric retrograde.

[0154] Step 202: Based on the B-plane parameter formula, determine the first parameter value of the P-plane parameter of the detector according to the vector parameter, and determine the second parameter value of the P-plane parameter according to the target parameter and the vector parameter; wherein, if the orbit to be determined is a hyperbolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the orbit asymptote, and if it is an elliptical orbit or a parabolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the pericenter vector;

[0155] As mentioned above, the definition of P-plane parameters is similar to that of B-plane parameters, except that the modulus of the B vector is replaced by the orbital semi-path p instead of the minor semi-axis b. For elliptical and parabolic orbits that are not applicable to the B-plane, Figure 1d The E vector shown in Figure 1c Thus, the parameter values ​​of the above-mentioned P-plane parameters can be determined based on the B-plane parameter formula.

[0156] When determining the first parameter value of the detector's P plane parameter according to the vector parameter in the above step 202, firstly, based on the B plane parameter formula, the n vector, the orbit semi-radius p and the E vector are determined according to the vector parameters; wherein the n vector is the orbit normal vector; and the E vector is the pericenter vector. Specifically, the n vector is determined according to the vector parameters by the following formula (14):

[0157]

[0158] Where n is the n vector, r and v are vector parameters, and μ is the Earth's gravitational constant. Then the orbit radius p is determined based on the vector parameters using the following formula (15):

[0159]

[0160] Where p is the semi-diameter of the track. Then the E vector is determined according to the vector parameters using the following formula (16):

[0161]

[0162] Further, based on the orbit type of the free return orbit, the P vector is determined according to the n vector, the p vector and the E vector. In implementation, if the orbit type is a hyperbolic orbit type, the S vector is determined according to the n vector and the E vector, and the P vector is determined according to the p vector, the S vector and the n vector. Specifically, the S vector is first determined according to the following formulas (17) to (18):

[0163] β=cos -1 (1 / e) Formula (17)

[0164]

[0165] Where β is the angle of the moon, S is the S vector, and E is the E vector. The S vector is a unit vector that passes through the center of the reference celestial body and is parallel to the asymptote of the hyperbolic orbit. Then, the P vector under the hyperbolic orbit type is determined according to the following formula (19):

[0166] P=p(S×n) Formula (19)

[0167] For the hyperbolic orbit type, the vector direction of the P vector is from the center of the reference celestial body to the intersection of the orbit asymptote and the P plane, and the magnitude of the P vector is the same as p.

[0168] If the orbit type is an elliptical orbit type or a parabolic orbit type, the P vector is determined according to the following formula (20);

[0169] P=p(E×n) Formula (20)

[0170] For an elliptical orbit type or a parabolic orbit type, the vector direction of the P vector is perpendicular to the E vector from the center of the reference celestial body and points to the side entering the orbit, and the magnitude of the P vector is the same as p.

[0171] Furthermore, for the hyperbolic track type, the T vector and the R vector are determined according to the S vector and the N vector by the following formulas (21) to (22):

[0172]

[0173] R=S×T Formula (22)

[0174] Among them, T and R are T vector and R vector respectively, T vector and R vector are unit vectors perpendicular to the P plane; N is a reference vector determined according to the rotation axis of the reference celestial body.

[0175] For an elliptical orbit type or a parabolic orbit type, the T vector and the R vector are determined from the E vector and the N vector by the following formulas (23) to (24):

[0176]

[0177] R=E×T Formula (24)

[0178] After the P vector is determined, the first parameter value is determined according to the T vector, the R vector and the P vector by the following formula (25):

[0179]

[0180] Among them, P T represents the projection of P vector under T vector, P R Represents the projection of the P vector under the R vector.

[0181] The first parameter value determined in the above formulas (14) to (25) does not take into account the task constraints in the actual engineering task. Next, substitute the target parameters that characterize the task constraints, and determine the second parameter value of the P-plane parameter based on the target parameters and the vector parameters. During implementation, first determine the semi-major axis a based on the vector parameters using the above formula (1), and then determine the angle α based on the orbit type of the free return orbit. It should be understood that in actual applications, the target P-plane parameters are often unknown. What is known are the target parameters such as orbital inclination and pericentricity, or re-entry point height and re-entry angle. The target P-plane parameters cannot be obtained directly, and other parameters are also needed. For orbits with eccentricities close to 1, such as the Earth-Moon transfer orbit or the Moon-Earth transfer orbit, the changes in S and E are very small and can be considered unchanged during iterative solutions. Therefore, the S vector, E vector and orbital semi-major axis p can be kept unchanged, that is, the S vector, E vector and p in the above formulas (14) to (25) are used, and then the second parameter value of the P-plane parameter is determined based on the target parameters and the vector parameters. As Figure 2e As shown, Figure 2e The angle α in is the angle between the P plane and the reference plane. The reference plane is a plane passing through the center of the reference celestial body and perpendicular to the n vector.

[0182] Specifically, if the orbit type is a hyperbolic type, the angle α is determined according to the component of the S vector in the z-axis direction of the preset inertial system by the following formula (26);

[0183] α=aeccos(S z ) Formula (26)

[0184] Among them, S z Characterizes the component of the S vector in the z-axis direction of the preset inertial system. If the orbit type is elliptical or parabolic, the angle α is determined according to the component of the E vector in the z-axis direction of the preset inertial system by the following formula (27):

[0185] α=aeccos(E z ) Formula (27)

[0186] Among them, E zCharacterizes the component of the E vector in the z-axis direction of the preset inertial system. After determining the angle α, for the hyperbolic track type, the track semi-diameter p is determined by the following formula (28):

[0187] p=r p (2-r p / a) Formula (28)

[0188] Among them, r p is the target pericentric distance in the target parameters.

[0189] For elliptical orbit type or parabolic orbit type, the orbit semi-diameter p is determined by the following formula (29):

[0190]

[0191] Among them, r e is the geocentric distance of the reentry point in the target parameters; γ is the reentry angle in the target parameters.

[0192] Based on the right spherical triangle formula, the angle is determined according to the target parameters and α using the following formula (30):

[0193]

[0194] Among them, the angle is the angle between the P vector and the T vector; i is the orbital inclination of the entry point in the target parameter. Finally, the following formula (31) is used to calculate the angle between p and the angle Determine the second parameter value:

[0195]

[0196] In actual engineering tasks, whether based on mission requirements or affected by perturbation factors, the free orbit around the moon cannot be located only in the ecliptic plane. Therefore, the space free return orbit around the moon is designed based on the P-plane parameters. The orbital force model can be a double two-body model or a high-precision force model.

[0197] In the instantaneous Earth-Moon inertial system, the starting point of the free return orbit around the Moon is selected in the xoz plane. At this time, y 0 =0, at the given starting point coordinate x 0 The design parameters of the latter orbit are [z0, vx, vy, vz] T. The orbit calculation method is the same as the design of the free return orbit around the moon in the ecliptic plane, that is, the entry parameters are calculated by reverse integration from the starting point, and the return parameters are calculated by forward integration. Then, the design variables are corrected according to the deviation between the two and the design target until the deviation meets the design requirements.

[0198] Step 203: Determine the iteration deviation of this iteration according to the first parameter value and the second parameter value; during implementation, after determining the first parameter value and the second parameter value of the P plane parameter through the above step 202, the difference between the first parameter value and the second parameter value is used as the iteration deviation of this iteration.

[0199] Step 204: For each iteration, if the iteration deviation is greater than a preset threshold, the initial value is corrected based on the differential correction method to obtain a corrected value of each track design parameter, and the corrected value is used as the initial value of each track design parameter in the next iteration, and the iteration process is repeated until the iteration deviation is no greater than the preset threshold.

[0200] The differential correction algorithm and B-plane parameters are widely used in deep space orbit design and are also the basic algorithms of this application. Assume that the constraint variable of the Earth-Moon transfer orbit is an m-dimensional vector P, and the control variable is an n-dimensional vector Q. The former can be determined according to the requirements of the target orbit of the lunar exploration mission; the latter takes the orbital parameters of the transfer orbit correction point, which is generally a velocity increment. There is a functional relationship between the two as shown in the following formula (32):

[0201] Q=f(P) Formula (32)

[0202] If the initial value of the control variable is P 0 , the corresponding constraint variable value is Q 0 , apply the above formula (32) to P 0 Perform Taylor expansion at and take the first-order term to obtain the following formula (33):

[0203]

[0204] Among them, M is the Jacobian matrix, which reflects the sensitivity of the constraint variables to small changes in the control variables. In engineering tasks, the difference method can be used for numerical solution. 0 With the target constraint Q s There is a deviation Q 0 -Q s , then the modified Q is iteratively calculated according to the following formula (34): 0 -Q s Required ΔP:

[0205] Δp=M -1 ΔQ formula (34)

[0206] Correspondingly, if m≠n, the following formula (35) can be used to calculate ΔP according to the minimum norm generalized inverse:

[0207] Δp=M T (M×M T ) -1 ΔQ formula (35)

[0208] The above process iterates the initial values ​​of each orbit design parameter based on the target parameters required by the mission constraints to determine the iterative deviation of the P-plane parameters of the probe. The orbit design parameters are corrected based on the iterative deviation to determine the free return orbit of different orbit configurations that meet the target parameters.

[0209] Next, the computational simulation process of the three-dimensional free return orbit is described to determine the feasibility of the technical solution of the present application.

[0210] For a given target inclination and perigee height, the P plane parameters have two states, corresponding to ascending and descending orbits. 0 , target entry inclination and perigee height, target return inclination and perigee height, there are 4 orbits. If analogous to the 4 configurations in the ecliptic plane, there are 16 configurations of the three-dimensional free return orbit around the moon. However, for the three-dimensional case, the geocentric prograde and retrograde are given by the target inclination, and there are 8 configurations for a given inclination and target perigee height. The following lunar position velocity at 2022-08-09T07:58:50 establishes the inertial system A, x 0 The free return orbit around the moon is calculated as ±6000 km. The inclination of the entry point is set to 28.5 degrees, the altitude is set to 200 kilometers, and the return point is set to an inclination of 43.0 degrees and an altitude of 120 kilometers.

[0211] Under the two-body model, the calculation results are as follows Figure 3a shown. Figure 3a The eight configurations under the double two-body model are shown (i.e. Figure 3a The free return orbit parameters of FRO1~8 in . Under the high-precision model, the calculation results are as follows Figure 3b shown. Figure 3b The free return orbit parameters of 8 configurations under the high-precision model are shown. Figure 3a and Figure 3b It can be seen that the orbits solved are all perigee orbital parameters, and the inclination is a set value, so the inclination and the horizontal anomaly are not listed. Comparing the double two-body model with the high-precision model, it can be found that the deviation of the double two-body model is not large, so the double two-body model can be used for preliminary window search and analysis. 0 6000 km, that is, the lunar retrograde orbit, the total flight time is about 6 days, and the x of the last 4 configurations 0 =-6000 km, i.e., the lunar center prograde orbit, and the total flight time is about 14 days. Therefore, if a manned lunar landing orbit is launched, a lunar center retrograde orbit should be adopted. Specifically, the three-dimensional view of the four configurations of lunar center retrograde free return orbit is as follows Figure 3c As shown, the four configurations are Figure 3cFRO1~4 shown in the figure; 4 configurations of free return orbits of the moon center prograde are shown in the figure. Figure 3d As shown, the four configurations are Figure 3d FRO5~8 shown in.

[0212] When the inclination of the geocentric orbit is greater than 90 degrees, it is a geocentric retrograde orbit. The situation is similar to this. The P-plane parameter method can quickly obtain an accurate orbit. 0 The transfer time can be adjusted, such as Figure 3e As shown, Figure 3e The same as the previous orbit entry point and return point settings are shown, but x 0 Total transfer time curves under different conditions. From the curves, we can see that the transfer time of the lunar center prograde orbit varies with |x 0 | is monotonically decreasing, and the lunar retrograde orbit is monotonically increasing. 0 The total transfer time is greatly affected, especially when the lunar center is at a low distance from the moon, the transfer time of the prograde orbit of the lunar center changes dramatically, while the transfer time of the retrograde orbit of the lunar center changes slowly. 0 At about 26,000 km, the transfer time of the lunar center sequence and the retrograde orbit is close, which confirms the effectiveness of the technical solution of this application.

[0213] Based on the same inventive concept, the embodiment of the present application provides a design device 400 for a detector free return track, such as Figure 4 As shown, including:

[0214] The parameter determination module 401 is configured to perform iterative processing on the initial values ​​of each track design parameter based on the target parameter, so as to modify each track design parameter according to the iterative deviation obtained in each iteration;

[0215] The orbit determination module 402 is configured to construct a free return orbit according to the orbit parameters that meet the engineering requirements after correction; wherein the orbit design parameters at least include the position component z of the probe on the z-axis of the preset inertial system. 0 and the velocity components v on the x, y, and z axes of the preset inertial system x 、v y 、v z The preset inertial system is a pre-constructed ecliptic plane symmetrical free return orbit instantaneous ecliptic coordinate system of the moon center at the perigee time; the target parameters include at least the first target parameter or the second target parameter; the first target parameter includes the orbital inclination i of the entry point and the target perigee distance r p The second target parameters include the reentry angle γ and the reentry point geocentric distance r e ;

[0216] The iterative processing process is as follows:

[0217] In the first iteration, the initial value is determined based on the pre-constructed symmetric free return orbit in the ecliptic plane, and the vector parameters corresponding to the initial value are determined by dynamic integration; wherein the vector parameters at least include the velocity vector v and the position vector r of the probe at the perigee or the velocity vector v and the position vector r of the probe at the re-entry point;

[0218] Based on the B-plane parameter formula, a first parameter value of the P-plane parameter of the detector is determined according to the vector parameter, and a second parameter value of the P-plane parameter is determined according to the target parameter and the vector parameter; wherein, if the orbit to be determined is a hyperbolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the orbit asymptote, and if it is an elliptical orbit or a parabolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the pericenter vector;

[0219] Determine an iteration deviation of this iteration according to the first parameter value and the second parameter value;

[0220] For each iterative processing, if the iterative deviation is greater than a preset threshold, the initial value is corrected based on the differential correction method to obtain the corrected value of each track design parameter, and the corrected value is used as the initial value of each track design parameter in the next iteration, and the iterative process is repeated until the iterative deviation is no greater than the preset threshold.

[0221] In some possible embodiments, the symmetrical free return orbit in the ecliptic plane is constructed with the perigee of the probe as the starting point; the starting point is located on the x-axis of the preset inertial system, and the preset inertial system is parallel to the instantaneous ecliptic coordinate system of the moon center at the perigee moment of the symmetrical free return orbit in the ecliptic plane; the earth-moon transfer orbit and the moon-earth return orbit of the symmetrical free return orbit in the ecliptic plane are symmetrical on both sides of the x-axis;

[0222] When performing the first iteration, the initial values ​​of the track design parameters are determined based on the pre-constructed symmetrical free return track in the white track plane, and the parameter determination module 401 is configured as follows:

[0223] Determine z according to the orbit structure of the symmetrical free return orbit in the ecliptic plane 0 、v x and v z The initial value of z 0 、v x and v z The value in the symmetric free return orbit model in the ecliptic plane is 0;

[0224] Based on the preset speed interval, the detector performs multiple acquisition operations on the parameters to be processed in the symmetrical free return orbit in the ecliptic plane; wherein the parameters to be processed at least include the position component x of the detector on the x-axis at each acquisition. 0 and the velocity component v on the y-axis 0 ;

[0225] After each acquisition, the pericentric distance r corresponding to the current acquisition operation is determined based on the parameters to be processed. pe , and according to the r corresponding to each acquisition operation pe Determine the v y The initial value of .

[0226] In some possible embodiments, the step of determining the pericentric distance r corresponding to the current round of acquisition operation based on the parameters to be processed is performed. pe , the parameter determination module 401 is configured as follows:

[0227] According to the x 0 and the v 0 Determine the semi-major axis a, the eccentricity e and the semi-diameter p of the orbit, and determine the true anomaly θ of the boundary point of the lunar influence sphere based on a, e and p;

[0228] Determine the radial velocity component v of the probe at the boundary of the lunar sphere of influence according to θ r , vertical radial velocity component v u and the flight time Δt, and according to the v r 、The v u and the Δt to determine the position vector R of the moon in the preset inertial system L and velocity vector V L ;

[0229] According to the R L and the V L Determine the position vector R of the detector at the center of the earth e and velocity vector V e , and according to the R e and the V e Determine the pericentric distance r corresponding to this round of acquisition operation pe .

[0230] In some possible embodiments, the B-plane parameter formula is executed to determine the first parameter value of the P-plane parameter of the detector according to the vector parameter, and the parameter determination module 401 is configured as follows:

[0231] Based on the B-plane parameter formula, the n vector, the orbit semi-radius p and the E vector are determined according to the vector parameters; wherein the n vector is the orbit normal vector; and the E vector is the pericenter vector;

[0232] Based on the orbit type of the free return orbit, according to the n vector, the p 2 And the E vector determines the P vector, and the first parameter value is determined according to the P vector.

[0233] In some possible embodiments, performing the step of: 2 And the E vector determines the P vector, the parameter determination module 401 is configured as follows:

[0234] If the orbit type is a hyperbolic orbit type, the S vector is determined according to the n vector and the E vector, and the P vector is determined according to the p, the S vector and the n vector; wherein, for the hyperbolic orbit type, the S vector is a unit vector passing through the center of the reference celestial body and having a direction parallel to the asymptote of the hyperbolic orbit, and the vector direction of the P vector is from the center of the reference celestial body to the intersection of the asymptote of the orbit and the P plane;

[0235] If the orbit type is the elliptical orbit type or the parabolic orbit type, the P vector is determined according to the p, the E vector and the n vector; wherein, for the elliptical orbit type or the parabolic orbit type, the vector direction of the P vector is perpendicular to the E vector from the center of the reference celestial body and points to the side entering the orbit.

[0236] In some possible embodiments, to determine the first parameter value according to the P vector, the parameter determination module 401 is configured to:

[0237] If the track type is a hyperbolic track type, a T vector and an R vector are determined according to the S vector and the N vector; wherein the T vector and the R vector are unit vectors perpendicular to the P plane;

[0238] If the orbit type is an elliptical orbit type or a parabolic orbit type, the T vector and the R vector are determined according to the E vector and the N vector; wherein the N vector is a reference vector determined according to the rotation axis of a reference celestial body.

[0239] The first parameter value is determined according to the T vector, the R vector, and the P vector.

[0240] In some possible embodiments, the determining of the second parameter value of the P-plane parameter according to the target parameter and the vector parameter is performed, and the parameter determination module 401 is configured as follows:

[0241] Determine the semi-major axis a according to the vector parameter;

[0242] Determine the angle α based on the orbit type of the free return orbit, and determine the orbit semi-path p according to the target parameter and a; wherein the angle α is the angle between the P plane and the reference plane; the reference plane is a plane passing through the center of the reference celestial body and perpendicular to the n vector;

[0243] Based on the right spherical triangle formula, the angle is determined according to the target parameter and α Among them, the angle is the angle between the P vector and the T vector;

[0244] According to the p and the angle The second parameter value is determined.

[0245] In some possible embodiments, the determination of the angle α based on the orbit type of the free return orbit is performed, and the parameter determination module 401 is configured as follows:

[0246] If the orbit type is the hyperbola type, determining the S vector according to the n vector and the E vector, and determining the angle α according to the component of the S vector in the z-axis direction of the preset inertial system;

[0247] If the orbit type is the ellipse type or the parabola type, the angle α is determined according to the component of the E vector in the z-axis direction of the preset inertial system.

[0248] Refer to the following Figure 5 The electronic device 130 according to this embodiment of the present application is described. Figure 5 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.

[0249] like Figure 5 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).

[0250] 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.

[0251] 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 .

[0252] 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.

[0253] 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.

[0254] 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.

[0255] 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 a probe free return trajectory as provided in the present application.

[0256] In an exemplary embodiment, various aspects of a method for designing a free return orbit of a probe provided in the present application may also be implemented in the form of a program product, which includes a 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 a free return orbit of a probe according to various exemplary embodiments of the present application described above in this specification.

[0257] 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.

[0258] 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.

[0259] 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.

[0260] 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.

[0261] 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).

[0262] 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.

[0263] 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.

[0264] 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.

[0265] 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 produce a device for implementing the functions specified in one or more processes in the flowchart and / or one or more boxes in the block diagram.

[0266] These computer program instructions may also be stored in a computer-readable memory that can direct 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 a product including an instruction device that implements the functions specified in one or more processes in the flowchart and / or one or more boxes in the block diagram.

[0267] 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 steps for implementing the functions specified in one or more flows in the flowchart and / or one or more blocks in the block diagram.

[0268] 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.

[0269] 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 design method for a probe's free return trajectory. It is characterized in that The method comprises: The initial values ​​of each orbital design parameter are iteratively processed based on the target parameter, so as to correct each orbital design parameter according to the iterative deviation obtained in each iteration, and a free return orbit is constructed according to each orbital parameter that meets the engineering requirements after correction; wherein the orbital design parameter at least includes the position component z of the probe on the z-axis of the preset inertial system 0 and the velocity components v on the x, y, and z axes of the preset inertial system x 、v y 、v z The preset inertial system is a pre-constructed ecliptic plane symmetrical free return orbit instantaneous ecliptic coordinate system of the moon center at the perigee time; the target parameters include at least the first target parameter or the second target parameter; the first target parameter includes the orbital inclination i of the entry point and the target perigee distance r p The second target parameters include the reentry angle γ and the reentry point geocentric distance r e ; The iterative processing process is as follows: In the first iteration, the initial value is determined based on the pre-constructed symmetric free return orbit in the ecliptic plane, and the vector parameters corresponding to the initial value are determined by dynamic integration; wherein the vector parameters at least include the velocity vector v and the position vector r of the probe at the perigee or the velocity vector v and the position vector r of the probe at the re-entry point; Based on the B-plane parameter formula, a first parameter value of the P-plane parameter of the detector is determined according to the vector parameter, and a second parameter value of the P-plane parameter is determined according to the target parameter and the vector parameter; wherein, if the orbit to be determined is a hyperbolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the orbit asymptote, and if it is an elliptical orbit or a parabolic orbit, the P-plane is a plane passing through the center of the reference celestial body and perpendicular to the pericenter vector; Determine an iteration deviation of this iteration according to the first parameter value and the second parameter value; For each iteration, if the iteration deviation is greater than a preset threshold, the initial value is corrected based on the differential correction method to obtain a corrected value of each track design parameter, and the corrected value is used as the initial value of each track design parameter in the next iteration, and the iteration process is repeated until the iteration deviation is no greater than the preset threshold; The determining the first parameter value of the P plane parameter of the detector according to the vector parameter based on the B plane parameter formula includes: determining the n vector, the orbit semi-radius p and the E vector according to the vector parameter based on the B plane parameter formula; wherein the n vector is the orbit normal vector; and the E vector is the pericenter vector; Based on the orbit type of the free return orbit, determine a P vector according to the n vector, the p vector, and the E vector, and determine the first parameter value according to the P vector; The determining the second parameter value of the P plane parameter according to the target parameter and the vector parameter comprises: determining the semi-major axis a according to the vector parameter; Determine the angle α based on the orbit type of the free return orbit, and determine the orbit semi-path p according to the target parameter and a; wherein the angle α is the angle between the P plane and the reference plane; the reference plane is a plane passing through the center of the reference celestial body and perpendicular to the n vector; Based on the right spherical triangle formula, the angle is determined according to the target parameter and α Among them, the angle is the angle between the P vector and the T vector; According to the p and the angle The second parameter value is determined.

2. The method according to claim 1, It is characterized in that The symmetrical free return orbit in the ecliptic plane is constructed with the perigee of the probe as the starting point; the starting point is located at the x-axis of the preset inertial system, and the preset inertial system is parallel to the instantaneous ecliptic coordinate system of the moon center at the perigee moment of the symmetrical free return orbit in the ecliptic plane; the earth-moon transfer orbit and the moon-earth return orbit of the symmetrical free return orbit in the ecliptic plane are symmetrical on both sides of the x-axis; In the first iteration, the initial values ​​of the orbit design parameters are determined based on the pre-constructed symmetrical free return orbit in the orbit plane, including: Determine z according to the orbit structure of the symmetrical free return orbit in the ecliptic plane 0 、v x and v z The initial value of z 0 、v x and v z The value in the symmetric free return orbit model in the ecliptic plane is 0; Based on the preset speed interval, the detector performs multiple acquisition operations on the parameters to be processed in the symmetrical free return orbit in the ecliptic plane; wherein the parameters to be processed at least include the position component x of the detector on the x-axis at each acquisition. 0 and the velocity component v on the y-axis 0 ; After each acquisition, the pericentric distance r corresponding to the current acquisition operation is determined based on the parameters to be processed. pe , and according to the r corresponding to each acquisition operation pe Determine the v y The initial value of .

3. The method according to claim 2, It is characterized in that The method of determining the pericentric distance r corresponding to the current round of acquisition operation based on the parameters to be processed pe ,include: According to the x 0 and the v 0 Determine the semi-major axis a, the eccentricity e and the semi-diameter p of the orbit, and determine the true anomaly θ of the boundary point of the lunar influence sphere based on a, e and p; Determine the radial velocity component v of the probe at the boundary of the lunar sphere of influence according to θ r , vertical radial velocity component v u and the flight time Δt, and according to the v r 、The v u and the Δt to determine the position vector R of the moon in the preset inertial system L and the velocity vector V L ; According to the R L and the V L Determine the position vector R of the detector at the center of the earth e and the velocity vector V e , and according to the R e and the V e Determine the pericentric distance r corresponding to this round of acquisition operation pe .

4. The method according to claim 1, It is characterized in that The determining the P vector according to the n vector, the p and the E vector comprises: If the orbit type is a hyperbolic orbit type, the S vector is determined according to the n vector and the E vector, and the P vector is determined according to the p, the S vector and the n vector; wherein, for the hyperbolic orbit type, the S vector is a unit vector passing through the center of the reference celestial body and having a direction parallel to the asymptote of the hyperbolic orbit, and the vector direction of the P vector is from the center of the reference celestial body to the intersection of the asymptote of the orbit and the P plane; If the orbit type is the elliptical orbit type or the parabolic orbit type, the P vector is determined according to the p, the E vector and the n vector; wherein, for the elliptical orbit type or the parabolic orbit type, the vector direction of the P vector is perpendicular to the E vector from the center of the reference celestial body and points to the side entering the orbit.

5. The method according to claim 4, It is characterized in that The determining the first parameter value according to the P vector includes: If the track type is a hyperbolic track type, a T vector and an R vector are determined according to the S vector and the N vector; wherein the T vector and the R vector are unit vectors perpendicular to the P plane; If the orbit type is an elliptical orbit type or a parabolic orbit type, the T vector and the R vector are determined according to the E vector and the N vector; wherein the N vector is a reference vector determined according to the rotation axis of a reference celestial body; The first parameter value is determined according to the T vector, the R vector, and the P vector.

6. The method according to claim 1, It is characterized in that The determining the angle α based on the track type of the free return track comprises: If the orbit type is a hyperbola type, determining the S vector according to the n vector and the E vector, and determining the angle α according to the component of the S vector in the z-axis direction of the preset inertial system; If the orbit type is an ellipse type or a parabola type, the angle α is determined according to the component of the E vector in the z-axis direction of the preset inertial system.

7. An electronic device, It is 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 1-6.

8. A computer storage medium, It is 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 6.

Images