CN115391722A - Space target acceleration display method and device, electronic equipment and storage medium - Google Patents
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Abstract
The invention provides a method, a device, electronic equipment and a storage medium for accelerated display of a space target. By adopting the method, the high-speed accelerated display of the mass space targets can be realized, so that the problem of orbit distortion in the conventional space target high-speed accelerated display scheme is solved.
Description
Technical Field
The invention relates to the technical field of space target operation, in particular to a space target accelerated display method, a space target accelerated display device, electronic equipment and a storage medium.
Background
Space targets include three types of spacecraft, rocket bodies, and space debris. The software displayed by the current spacecraft is only used for panoramic display of the spacecraft, but the space target is not only used by the spacecraft, but also comprises a large amount of rocket debris and space debris.
In the process of implementing the invention, the inventor finds that a stuck phenomenon occurs when the current scheme is adopted to load the operation scene of the whole space target, so that the panoramic display effect of the whole space target is poor. In the related technology, a Lagrange interpolation method can be used for solving the problem of space target display blockage, but when high-speed acceleration display is carried out, along with the increase of time span, a satellite operation orbit cannot be approximately fitted by using a polynomial, so that the Lagrange interpolation method is not applicable any more, and the space target orbit can deviate from the original orbit to cause orbit distortion.
Disclosure of Invention
In view of the above, the present invention provides a method, an apparatus, an electronic device and a storage medium for displaying space targets in an accelerated manner, so as to alleviate the above problems in the related art.
In a first aspect, an embodiment of the present invention provides a method for accelerated display of a space target, where the method includes: acquiring initial orbit root data of a space target in a preset time period; the initial orbit root data comprise orbit parameters of the space target and a mean-near point angle corresponding to the initial moment of the space target in a preset time period; dividing the preset time interval into a plurality of target moments with continuous time according to a preset speed; wherein the first target time is after the initial time; calculating to obtain the position coordinates of the space target at each target moment under a preset celestial body geodetic coordinate system according to the initial orbit root data; and rendering to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment in a preset celestial body geodetic coordinate system.
As one possible implementation, the orbit parameters include period, eccentricity, semi-major axis, orbit inclination, ascension at the point of ascension, and argument of perigee; according to the initial orbit root data, calculating to obtain the position coordinates of the space target at each target moment under a preset celestial body geodetic coordinate system, wherein the step comprises the following steps of: calculating to obtain a flat near point angle corresponding to the space target at each target time according to the operation period of the space target and the flat near point angle corresponding to the space target at the initial time; and for each target moment, calculating to obtain the position coordinates of the space target at the target moment under a preset celestial body geodetic coordinate system according to the eccentricity, the semimajor axis, the orbit inclination angle, the ascension of the ascending intersection point and the corresponding approximate point angle of the space target at the target moment.
As a possible implementation, the step of calculating the position coordinate of the space target at the target time under a preset celestial body geodetic coordinate system according to the eccentricity, the semimajor axis, the orbit inclination angle, the ascension of the ascending intersection point and the corresponding approximate point angle of the space target at the target time comprises: calculating to obtain the position coordinates of the space target at the target moment under a preset orbit coordinate system according to the eccentricity and the semimajor axis of the space target and the mean-near point angle corresponding to the space target at the target moment; and converting the position coordinates of the target time space target under a preset orbit coordinate system into the position coordinates of the target time space target under a preset celestial body and terrestrial coordinate system according to the semimajor axis, the orbit inclination, the ascension point and the perigee argument of the space target.
As a possible implementation, the step of obtaining the moving track of the space target within a preset time period by rendering according to the position coordinate of the space target at each target moment under a preset celestial body geodetic coordinate system comprises: and rendering in a multithread rendering mode to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment in a preset celestial body geodetic coordinate system.
In a second aspect, an embodiment of the present invention further provides a space target accelerated display device, where the device includes: the acquisition module is used for acquiring initial orbit root data of the space target in a preset time period; the initial orbit root data comprise orbit parameters of the space target and a mean-near point angle corresponding to the initial moment of the space target in a preset time period; the dividing module is used for dividing the preset time period into a plurality of target moments with continuous time according to a preset speed; wherein the first target time is after the initial time; the calculation module is used for calculating and obtaining the position coordinates of the space target under a preset celestial body geodetic coordinate system at each target moment according to the initial orbit root data; and the rendering module is used for rendering to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment under a preset celestial body geodetic coordinate system.
As one possible implementation, the orbit parameters include period, eccentricity, semi-major axis, orbit inclination, ascension at the point of ascension, and argument of perigee; the calculation module is further to: calculating to obtain a flat near point angle corresponding to the space target at each target time according to the operation period of the space target and the flat near point angle corresponding to the space target at the initial time; and for each target moment, calculating to obtain the position coordinates of the space target at the target moment under a preset celestial body geodetic coordinate system according to the eccentricity, the semimajor axis, the orbit inclination angle, the ascension of the ascending intersection point and the corresponding approximate point angle of the space target at the target moment.
As one possible implementation, the computing module is further configured to: calculating to obtain the position coordinates of the space target at the target moment under a preset orbit coordinate system according to the eccentricity and the semimajor axis of the space target and the mean-near point angle corresponding to the space target at the target moment; and converting the position coordinates of the target time space target under a preset orbit coordinate system into the position coordinates of the target time space target under a preset celestial body and terrestrial coordinate system according to the semimajor axis, the orbit inclination, the ascension point and the perigee argument of the space target.
As one possible implementation, the rendering module is further configured to: and rendering in a multi-thread rendering mode to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment in a preset celestial body and earth coordinate system.
In a third aspect, an embodiment of the present invention further provides an electronic device, including a processor and a memory, where the memory stores computer-executable instructions that can be executed by the processor, and the processor executes the computer-executable instructions to implement the foregoing method.
In a fourth aspect, embodiments of the present invention also provide a computer-readable storage medium storing computer-executable instructions that, when invoked and executed by a processor, cause the processor to implement the above-mentioned method.
The embodiment of the invention provides a space target acceleration display method, a space target acceleration display device, electronic equipment and a storage medium, wherein initial orbit root data of a space target in a preset time period are obtained firstly, and the initial orbit root data comprise orbit parameters of the space target and a mean-near point angle corresponding to an initial moment of the space target in the preset time period; dividing a preset time interval into a plurality of target moments with continuous time according to a preset speed; and then calculating to obtain the position coordinates of the space targets at each target moment under a preset celestial body and geodetic coordinate system according to the initial orbit root data, and finally rendering to obtain the running tracks of the space targets in a preset time interval according to the position coordinates of the space targets at each target moment under the preset celestial body and geodetic coordinate system. By adopting the technology, the high-speed accelerated display of mass space targets can be realized, so that the problem of orbit distortion in the conventional space target high-speed accelerated display scheme is solved.
Additional features and advantages of the invention will be set forth in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. The objectives and other advantages of the invention will be realized and attained by the structure particularly pointed out in the written description and claims hereof as well as the appended drawings.
In order to make the aforementioned and other objects, features and advantages of the present invention comprehensible, preferred embodiments accompanied with figures are described in detail below.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts.
FIG. 1 is a schematic flow chart of a space target accelerated display method according to an embodiment of the present invention;
FIG. 2 is a schematic structural diagram of a space target acceleration display device according to an embodiment of the present invention;
fig. 3 is a schematic structural diagram of an electronic device according to an embodiment of the present disclosure.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions of the present invention will be clearly and completely described below with reference to the embodiments, and it is obvious that the described embodiments are some, but not all embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
Space targets include three types of spacecraft, rocket bodies, and space debris. The software displayed by the current spacecraft is good in panoramic display effect only for the spacecraft, but the space target is not only the spacecraft, but also comprises a large amount of rocket debris and space debris.
In the process of implementing the invention, the inventor finds that a stuck phenomenon occurs when the current scheme is adopted to load the operation scene of the whole space target, so that the panoramic display effect of the whole space target is poor. In the related technology, a Lagrange interpolation method can be used for solving the problem of space target display blockage, but when high-speed acceleration display is carried out, along with the increase of time span, a satellite operation orbit cannot be approximately fitted by using a polynomial, so that the Lagrange interpolation method is not applicable any more, and the space target orbit can deviate from the original orbit to cause orbit distortion.
Based on the above, the space target accelerated display method, device, electronic device and storage medium provided by the implementation of the invention can alleviate the above problems in the related art.
To facilitate understanding of the embodiment, first, a detailed description is given to a space target acceleration display method disclosed in the embodiment of the present invention, referring to a flow diagram of a space target acceleration display method shown in fig. 1, where the method may include the following steps:
step S102, acquiring initial orbit number data of a space target in a preset time period; the initial orbit number data comprise orbit parameters of the space target and a mean-near point angle corresponding to the initial moment of the space target in a preset time period.
The space targets may include, but are not limited to, spacecraft, rocket bodies, space debris, and the like. The preset time interval can be a time interval in which the operation situation (namely, the operation track) of the space target needs to be displayed, and can be determined according to actual requirements, and the time interval is not limited.
Step S104, dividing a preset time interval into a plurality of target moments with continuous time according to a preset multiple speed; wherein the first target time may be after the initial time.
The preset multiple speed can be a display multiple required when the space target running track is displayed in an accelerated mode, and can be determined according to actual requirements, and the display multiple is not limited. Accordingly, the number of the target moments divided by the preset time period and the time interval between two adjacent target moments can be adjusted according to the preset multiple, which is not limited.
And S106, calculating to obtain the position coordinates of the space target under a preset celestial body geodetic coordinate system at each target moment according to the initial orbit root data.
The preset celestial body geodetic coordinate system is usually a three-dimensional spherical coordinate system with a certain celestial body as a coordinate origin, and the position coordinates under the coordinate system can be used for representing the position of the space target relative to the center of the celestial body.
And S108, according to the position coordinates of the space target at each target moment in a preset celestial body geodetic coordinate system, rendering to obtain the running track of the space target in a preset time period.
The space target accelerated display method provided by the embodiment of the invention comprises the steps of firstly obtaining initial orbit radical data of a space target in a preset time period, wherein the initial orbit radical data comprises orbit parameters of the space target and a mean-anomaly angle corresponding to the initial moment of the space target in the preset time period, and the orbit parameters comprise a period, an eccentricity ratio, a semi-major axis, an orbit inclination angle, a rising intersection right ascension and a near-anomaly argument; dividing a preset time interval into a plurality of target moments with continuous time according to a preset speed; and finally, according to the position coordinates of the space targets at each target moment in the preset celestial body-terrestrial coordinate system, rendering to obtain the running track of the space targets in a preset time period. By adopting the technology, the high-speed acceleration display of massive space targets can be realized, so that the problem of track distortion in the existing space target high-speed acceleration display scheme is solved.
As a possible implementation, the track parameters may include period, eccentricity, semi-major axis, track inclination, ascension at the intersection point, and argument of perigee; the initial orbit root data may be specifically an orbit root of a space target in a satellite orbit, which is solved by using a Two-Line Element (TLE file) corresponding to an initial time of the space target in a preset time period based on a Simplified General Perturbations (SGP 4) model, and is a period, an eccentricity, a semi-major axis, an orbit inclination, a rising point right ascension, a perigee argument and a perigee angle. Based on the method, a TLE file corresponding to the space target at the initial moment of a preset time period can be obtained, and then the number of orbit elements (including period, eccentricity, semimajor axis, orbit inclination angle, ascent point right-ascent, argument of near place and argument of near point) of the space target in the satellite orbit is solved by utilizing the TLE file based on an SGP4 model, so that the orbit elements are obtainedAnd taking the orbit number of the space target in the satellite orbit obtained by the solution as initial orbit number data of the space target in a preset time period. In the specific solving process, the mean-near-point angle corresponding to the space target at the initial moment can be calculated according to the following formula: m 0 = n (t 0 - t F0 ) (ii) a Wherein, M 0 Is the mean anomaly angle, t, corresponding to the space target at the initial moment F0 The time when the space target passes through the near place of the celestial body is taken as the time; t is t 0 The initial time is the initial time; n is the average angular velocity of the spatial object, in rad/s,
(ii) a a is the major radius of the orbital ellipse (i.e., the semi-major axis), G is the gravitational constant, and M e The celestial mass. The celestial body may be, but is not limited to, the earth or other known celestial bodies.
As a possible implementation manner, the step S106 (i.e. calculating the position coordinates of the space object under the preset celestial body and terrestrial coordinate system at each target time according to the initial orbit number data) may include the following operation manners:
(11) And calculating to obtain the mean-near point angle corresponding to the space target at each target time according to the operation cycle of the space target and the mean-near point angle corresponding to the space target at the initial time.
(12) And for each target moment, calculating to obtain the position coordinates of the space target at the target moment under a preset celestial body geodetic coordinate system according to the eccentricity, the semimajor axis, the orbit inclination angle, the ascension of the ascending intersection point and the corresponding approximate point angle of the space target at the target moment.
Exemplarily, the operation mode of the above (12) may specifically be: for each target moment, calculating to obtain the position coordinates of the target moment space target under a preset orbit coordinate system according to the eccentricity and the semimajor axis of the space target and the mean anomaly angle corresponding to the space target at the target moment, and then converting the position coordinates of the target moment space target under the preset orbit coordinate system into the position coordinates of the target moment space target under the preset celestial body geodetic coordinate system according to the semimajor axis, the orbit inclination angle, the ascent point right ascension and the anomaly argument of the space target.
For ease of understanding, the operation of the above (11) and (12) is described herein as follows, by way of example, with one specific example:
firstly, establishing an orbit coordinate system, wherein the coordinate origin of the coordinate system is positioned at the center of a celestial body, the X-axis and the Y-axis of the orbit coordinate system are positioned on an orbit plane, the Z-axis of the orbit coordinate system is superposed with the normal vector of the orbit plane, and the orbit coordinate system is a right-hand coordinate system;
for each target time, the mean-near point angle corresponding to the space target at the target time can be calculated according to the following formula according to the operation cycle of the space target and the mean-near point angle corresponding to the space target at the initial time: m k = M 0 + (t k - t 0 ) X 2 × pi/(T × 60 × 1000); wherein, M k Mean angle of approach, M, for space target at kth target time 0 Is the mean anomaly angle, t, corresponding to the space target at the initial moment k Is the kth target time, t 0 Taking k as an integer greater than 0, wherein T is the operation period of the space target at the initial moment;
for each target moment, after the mean anomaly corresponding to the space target at the target moment is obtained through calculation, the Kepler equation E can be solved according to the eccentricity of the space target and the mean anomaly corresponding to the space target at the target moment k =M k +e·sinE k Calculating to obtain the approximate point angle, E, corresponding to the space target at the target time k Is the approximate point angle, M, corresponding to the k-th target moment of the space target k A mean-near point angle corresponding to the space target at the kth target moment is obtained, and e is the eccentricity; adopting an angle system during resolving, and combining M k Substituting e into kepler's equation, and iterating until convergence (i.e.
And epsilon is a preset convergence threshold value,
is E obtained after the j +1 th iteration k Value of,
is E obtained after the jth iteration k Value) is stopped, so that the E obtained when stopping the iteration k The value is used as a near point angle corresponding to the k target time of the space target;
for each target time, after the approximate point angle corresponding to the space target at the target time is obtained through calculation, the distance between the space target and the celestial body center at the target time can be obtained through calculation according to the following formula according to the eccentricity and the semi-major axis of the space target and the approximate point angle corresponding to the space target at the target time:
(ii) a Wherein r is k Is the distance between the space target and the celestial body center at the kth target moment, a is the semimajor axis, E is the eccentricity, E is the distance between the space target and the celestial body center k A near point angle corresponding to the k target moment of the space target;
for each target time, after the distance between the space target and the celestial body center at the target time is obtained through calculation, according to the distance between the space target and the celestial body center at the target time and the approximate point angle corresponding to the space target at the target time, the true approximate point angle corresponding to the space target at the target time can be obtained through calculation according to the following formula:
(ii) a Wherein, theta is the true near point angle corresponding to the k-th target moment of the space target, E is the eccentricity, E is the distance between the k-th target moment and the true near point angle k A deviation from a near point angle corresponding to the space target at the kth target moment;
for each target time, after the true near point angle corresponding to the space target at the target time is obtained through calculation, the distance rate between the space target and the celestial body center and the space time can be used according to the target timeCalculating the position coordinate of the space target at the target moment under a preset orbit coordinate system according to the following formula: x k "=r k cosθ，Y k "=r k sinθ，Z k "=0; wherein r is k Is the distance between the space target and the celestial body center at the kth target time, theta is the true near point angle corresponding to the space target at the kth target time, X k "、Y k "and Z k "are respectively coordinate values corresponding to the space target at the kth target moment in three coordinate axis directions under a preset orbit coordinate system;
the orbit coordinate system can be coincided with the celestial body earth center rectangular coordinate system (namely the preset celestial body earth coordinate system) only through three rotations, and the coordinate axes of the orbit coordinate system are respectively an X-axis, a Y-axis and a Z-axis, and the coordinate axes of the celestial body earth center rectangular coordinate system are respectively an X-axis, a Y-axis and a Z-axis
A shaft,
Shaft and
the rotation process of the axis, orbital coordinate system may include: firstly, the Z 'axis of the orbit coordinate system is rotated counterclockwise by an angle of omega to lead the X' axis of the orbit coordinate system to point to the ascending point, and then the Z 'axis of the orbit coordinate system is rotated counterclockwise by an angle of i to lead the Z' axis of the orbit coordinate system and the celestial body geocentric rectangular coordinate system
With coincident axes and finally around the earth's centre rectangular coordinate system
The shaft rotates counterclockwise by one (omega-
) Angle of (2)So that the two coordinate systems coincide; wherein, omega is the amplitude angle of the near place, i is the track inclination angle, omega is the ascension of the ascending intersection point, and a is the semimajor axis;
is a rectangular coordinate system of the earth center of the celestial body
The included angle between the axis and the X' axis direction of the spring equinox point is the Greenwich mean hour angle; then, according to the semimajor axis, the orbit inclination angle, the ascension point right ascension and the perigee argument of the space target, a matrix R1, a matrix R2 and a matrix R3 are respectively constructed according to the following three formulas:
for each target moment, after the position coordinates of the target moment space target under the preset orbital coordinate system are obtained through calculation, the position coordinates of the target moment space target under the preset orbital coordinate system can be converted into the position coordinates of the target moment space target under the preset celestial body geodetic coordinate system according to the following formula:
(ii) a Wherein the content of the first and second substances,
the position coordinates of the space target at the kth target moment under a preset celestial body geodetic coordinate system,
and the position coordinates of the space target at the kth target moment under a preset orbit coordinate system.
As another specific example, for each target time, after the approximate point angle corresponding to the space target at the target time is obtained through calculation, the position coordinate of the space target at the target time in the preset orbital coordinate system at the target time may be directly obtained through calculation according to the following formula, based on the eccentricity and the semi-major axis of the space target and the approximate point angle corresponding to the space target at the target time:
，
，Z k "=0; wherein a is a semimajor axis, E is an eccentricity, E k A near point angle X corresponding to the k-th target time of the space target k "、Y k "and Z k "are coordinate values respectively corresponding to the k-th target time space target in three coordinate axis directions under a preset orbit coordinate system, and k is an integer larger than 0.
As a possible implementation manner, step S108 (i.e. rendering the trajectory of the space object within the preset time period according to the position coordinates of the space object under the preset celestial body and terrestrial coordinate system at each target time) may include the following operation manners: and rendering in a multi-thread rendering mode to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment in a preset celestial body and earth coordinate system.
Exemplarily, after the position coordinates of the space target under the preset celestial body geodetic coordinate system at each target moment in the preset time period are obtained through calculation, the calculation and rendering can be separated through web worker multithreading in JavaScript, and therefore the running track of the space target in the preset time period is obtained through rendering.
Based on the space target accelerated display method, an embodiment of the present invention further provides a space target accelerated display device, as shown in fig. 2, the device may include the following modules:
the acquisition module 202 is used for acquiring initial orbit number data of the space target in a preset time period; the initial orbit root data comprise orbit parameters of the space target and a mean-near point angle corresponding to the initial moment of the space target in a preset time period.
The dividing module 204 is configured to divide the preset time period into a plurality of target moments with continuous time according to a preset multiple speed; wherein the first target time is after the initial time.
And the calculating module 206 is configured to calculate and obtain a position coordinate of the space target at each target time in a preset celestial body geodetic coordinate system according to the initial orbit root data.
And the rendering module 208 is configured to render to obtain a running track of the space target within a preset time period according to the position coordinate of the space target at each target moment in a preset celestial body geodetic coordinate system.
The space target acceleration display device provided by the embodiment of the invention comprises the following steps of firstly obtaining initial orbit root data of a space target in a preset time period, wherein the initial orbit root data comprises orbit parameters of the space target and a mean-near point angle corresponding to an initial moment of the space target in the preset time period; dividing a preset time interval into a plurality of target moments with continuous time according to a preset speed; and finally, according to the position coordinates of the space targets at each target moment in the preset celestial body-terrestrial coordinate system, rendering to obtain the running track of the space targets in a preset time period. By adopting the technology, the high-speed accelerated display of mass space targets can be realized, so that the problem of orbit distortion in the conventional space target high-speed accelerated display scheme is solved.
The track parameters can comprise period, eccentricity, semi-major axis, track inclination angle, ascent point right ascension and perigee amplitude angle; based on this, the calculating module 206 may further be configured to: calculating to obtain a flat near point angle corresponding to the space target at each target time according to the operation period of the space target and the flat near point angle corresponding to the space target at the initial time; and for each target moment, calculating to obtain the position coordinates of the space target at the target moment under a preset celestial body geodetic coordinate system according to the eccentricity, the semimajor axis, the orbit inclination angle, the ascension of the ascending intersection point and the corresponding approximate point angle of the space target at the target moment.
The calculating module 206 may further be configured to: calculating to obtain the position coordinates of the space target at the target moment under a preset orbit coordinate system according to the eccentricity and the semimajor axis of the space target and the mean-near point angle corresponding to the space target at the target moment; and converting the position coordinates of the target time space target under a preset orbit coordinate system into the position coordinates of the target time space target under a preset celestial body and geodetic coordinate system according to the semimajor axis, the orbit inclination angle, the ascension point right ascension and the perigee argument of the space target.
The rendering module 208 may further be configured to: and rendering in a multi-thread rendering mode to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment in a preset celestial body and earth coordinate system.
The implementation principle and the generated technical effect of the space target acceleration display device provided by the embodiment of the invention are the same as those of the space target acceleration display method embodiment, and for brief description, the corresponding contents in the space target acceleration display method embodiment can be referred to where the device embodiment is not mentioned.
An embodiment of the present invention further provides an electronic device, as shown in fig. 3, which is a schematic structural diagram of the electronic device, where the electronic device includes a processor 31 and a memory 30, the memory 30 stores computer-executable instructions that can be executed by the processor 31, and the processor 31 executes the computer-executable instructions to implement the space target accelerated display method.
In the embodiment shown in fig. 3, the electronic device further comprises a bus 32 and a communication interface 33, wherein the processor 31, the communication interface 33 and the memory 30 are connected by the bus 32.
The Memory 30 may include a high-speed Random Access Memory (RAM) and may also include a non-volatile Memory (non-volatile Memory), such as at least one disk Memory. The communication connection between the network element of the system and at least one other network element is realized through at least one communication interface 33 (which may be wired or wireless), and the internet, a wide area network, a local network, a metropolitan area network, and the like may be used. The bus 32 may be an ISA (Industry Standard Architecture) bus, a PCI (Peripheral Component Interconnect) bus, an EISA (Extended Industry Standard Architecture) bus, or the like. The bus 32 may be divided into an address bus, a data bus, a control bus, etc. For ease of illustration, only one double-headed arrow is shown in FIG. 3, but this does not indicate only one bus or one type of bus.
The processor 31 may be an integrated circuit chip having signal processing capabilities. In implementation, the steps of the above method may be performed by integrated logic circuits of hardware or instructions in the form of software in the processor 31. The Processor 31 may be a general-purpose Processor, and includes a Central Processing Unit (CPU), a Network Processor (NP), and the like; the device can also be a Digital Signal Processor (DSP), an Application Specific Integrated Circuit (ASIC), a Field Programmable Gate Array (FPGA) or other Programmable logic device, a discrete Gate or transistor logic device, or a discrete hardware component. A general purpose processor may be a microprocessor or the processor may be any conventional processor or the like. The steps of the method disclosed in connection with the embodiments of the present invention may be directly implemented by a hardware decoding processor, or implemented by a combination of hardware and software modules in the decoding processor. The software module may be located in ram, flash memory, rom, prom, or eprom, registers, etc. storage media as is well known in the art. The storage medium is located in the memory, and the processor 31 reads the information in the memory, and completes the steps of the space object accelerated display method of the foregoing embodiment in combination with the hardware thereof.
The embodiment of the present invention further provides a computer-readable storage medium, where the computer-readable storage medium stores computer-executable instructions, and when the computer-executable instructions are called and executed by a processor, the computer-executable instructions cause the processor to implement the space target accelerated display method, and specific implementation may refer to the foregoing method embodiment, and is not described herein again.
The space target accelerated display method, the space target accelerated display device and the computer program product of the electronic device provided by the embodiment of the invention comprise a computer readable storage medium storing program codes, instructions included in the program codes can be used for executing the method described in the previous method embodiment, and specific implementation can refer to the method embodiment, and is not described herein again.
Unless specifically stated otherwise, the relative steps, numerical expressions, and values of the components and steps set forth in these embodiments do not limit the scope of the present invention.
The functions, if implemented in the form of software functional units and sold or used as a stand-alone product, may be stored in a non-volatile computer-readable storage medium executable by a processor. Based on such understanding, the technical solution of the present invention may be embodied in the form of a software product, which is stored in a storage medium and includes instructions for causing a computer device (which may be a personal computer, a server, or a network device) to execute all or part of the steps of the method according to the embodiments of the present invention. And the aforementioned storage medium includes: a U-disk, a removable hard disk, a Read-Only Memory (ROM), a Random Access Memory (RAM), a magnetic disk or an optical disk, and other various media capable of storing program codes.
In the description of the present invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, and are only for convenience of description and simplicity of description, but do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be construed as limiting the present invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
Finally, it should be noted that: the above-mentioned embodiments are only specific embodiments of the present invention, which are used for illustrating the technical solutions of the present invention and not for limiting the same, and the protection scope of the present invention is not limited thereto, although the present invention is described in detail with reference to the foregoing embodiments, those skilled in the art should understand that: any person skilled in the art can modify or easily conceive the technical solutions described in the foregoing embodiments or equivalent substitutes for some technical features within the technical scope of the present disclosure; such modifications, changes or substitutions do not depart from the spirit and scope of the embodiments of the present invention, and they should be construed as being included therein. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.
Claims (10)
1. A space target accelerated display method is characterized by comprising the following steps:
acquiring initial orbit root data of a space target in a preset time period; the initial orbit root data comprise orbit parameters of the space target and a mean-near point angle corresponding to the initial moment of the space target in a preset time period;
dividing the preset time interval into a plurality of target moments with continuous time according to a preset speed; wherein the first target time is after the initial time;
calculating to obtain the position coordinates of the space target at each target moment in a preset celestial body geodetic coordinate system according to the initial orbit root data;
and rendering to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment in a preset celestial body geodetic coordinate system.
2. The method of claim 1, wherein the orbital parameters include period, eccentricity, semi-major axis, orbital inclination, ascent point right ascension, and perigee argument; according to the initial orbit root data, calculating to obtain the position coordinates of the space target at each target moment under a preset celestial body geodetic coordinate system, wherein the step comprises the following steps of:
calculating to obtain a flat near point angle corresponding to the space target at each target time according to the operation period of the space target and the flat near point angle corresponding to the space target at the initial time;
and for each target moment, calculating to obtain the position coordinates of the space target at the target moment under a preset celestial body geodetic coordinate system according to the eccentricity, the semimajor axis, the orbit inclination angle, the ascension of the ascending intersection point and the corresponding approximate point angle of the space target at the target moment.
3. The method as claimed in claim 2, wherein the step of calculating the position coordinates of the space target at the target time under the preset celestial body and terrestrial coordinate system according to the eccentricity, semi-major axis, orbit inclination angle and ascent crossing right ascension of the space target and the mean-near point angle of the space target at the target time comprises:
calculating to obtain the position coordinates of the space target at the target moment under a preset orbit coordinate system according to the eccentricity and the semimajor axis of the space target and the mean-near point angle corresponding to the space target at the target moment;
and converting the position coordinates of the target time space target under a preset orbit coordinate system into the position coordinates of the target time space target under a preset celestial body and terrestrial coordinate system according to the semimajor axis, the orbit inclination, the ascension point and the perigee argument of the space target.
4. The method as claimed in claim 1, wherein the step of rendering the trajectory of the space target in the preset time period according to the position coordinates of the space target in the preset celestial body and terrestrial coordinate system at each target moment comprises:
and rendering in a multi-thread rendering mode to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment in a preset celestial body and earth coordinate system.
5. A space target accelerated display device, characterized in that the device comprises:
the acquisition module is used for acquiring initial orbit root data of the space target in a preset time period; the initial orbit root data comprise orbit parameters of the space target and a mean-near point angle corresponding to the initial moment of the space target in a preset time period;
the dividing module is used for dividing the preset time period into a plurality of target moments with continuous time according to a preset speed; wherein the first target time is after the initial time;
the calculation module is used for calculating and obtaining the position coordinates of the space target under a preset celestial body geodetic coordinate system at each target moment according to the initial orbit root data;
and the rendering module is used for rendering to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment under a preset celestial body geodetic coordinate system.
6. The apparatus of claim 5, wherein the orbital parameters include period, eccentricity, semi-major axis, orbital inclination, ascension at ascending intersection, and argument of perigee; the calculation module is further to:
calculating to obtain a flat near point angle corresponding to the space target at each target time according to the operation period of the space target and the flat near point angle corresponding to the space target at the initial time;
and for each target moment, calculating to obtain the position coordinates of the space target at the target moment under a preset celestial body and geodetic coordinate system according to the eccentricity, the semi-major axis, the orbit inclination angle, the ascent point and the ascent point of the space target and the approximate point angle corresponding to the space target at the target moment.
7. The apparatus of claim 6, wherein the computing module is further configured to:
calculating to obtain the position coordinates of the space target at the target moment under a preset orbit coordinate system according to the eccentricity and the semimajor axis of the space target and the mean-near point angle corresponding to the space target at the target moment;
and converting the position coordinates of the target time space target under a preset orbit coordinate system into the position coordinates of the target time space target under a preset celestial body and terrestrial coordinate system according to the semimajor axis, the orbit inclination, the ascension point and the perigee argument of the space target.
8. The apparatus of claim 5, wherein the rendering module is further configured to:
and rendering in a multi-thread rendering mode to obtain the running track of the space target in a preset time period according to the position coordinate of the space target at each target moment in a preset celestial body and earth coordinate system.
9. An electronic device comprising a processor and a memory, the memory storing computer-executable instructions executable by the processor, the processor executing the computer-executable instructions to implement the method of any of claims 1 to 4.
10. A computer-readable storage medium having computer-executable instructions stored thereon which, when invoked and executed by a processor, cause the processor to perform the method of any of claims 1 to 4.
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