CN115258198A - Spacecraft orbit determination method and device, processor and electronic equipment - Google Patents
Spacecraft orbit determination method and device, processor and electronic equipment Download PDFInfo
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Abstract
The application discloses a method and a device for determining a spacecraft orbit, a processor and electronic equipment. The method comprises the following steps: constructing a rotation matrix according to a target attitude quaternion corresponding to the target attitude of the target spacecraft; constructing a first partial derivative model of the unloading acceleration under the inertial coordinate system to the acceleration under the body coordinate system according to the rotation matrix, and selecting a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic; according to the second partial derivative model and observation data of the target spacecraft, a state equation of the target spacecraft is constructed, and integral solution is carried out to obtain target acceleration; and determining the orbit of the target spacecraft according to the target acceleration. By the method and the device, the problem that acceleration generated by unloading of the spacecraft momentum wheel in the related technology is mainly obtained by calculation through a uniform acceleration model under an orbit coordinate system or a celestial sphere reference system, so that the unloading acceleration is not stably solved, and further the orbit determination error is large is solved.
Description
Technical Field
The application relates to the technical field of spacecraft engineering, in particular to a method and a device for determining a spacecraft orbit, a processor and electronic equipment.
Background
The momentum wheel mainly utilizes the principle of acting force and reacting force, when external interference force acts on the spacecraft to generate interference torque, the momentum wheel arranged at the corresponding position of the spacecraft resists the influence of the interference torque by increasing the rotating speed and increasing the momentum. When the rotating speed of the momentum wheel reaches a threshold value, the disturbance torque can be counteracted by air injection, and meanwhile, the accumulated momentum is released.
Momentum wheel unloading is an important factor influencing the resolving accuracy of the detector orbit, and the acceleration generated by momentum wheel unloading is mainly solved by adopting a uniform acceleration model under an orbit coordinate system or a celestial sphere reference system, but the acceleration is easy to be inaccurately calculated under the condition of poor observation geometry by the model.
Aiming at the problems that in the related art, the acceleration generated by unloading the spacecraft momentum wheel is mainly obtained by calculating a uniform acceleration model under an orbit coordinate system or a celestial sphere reference system, so that the unloading acceleration is not stably solved, and further the orbit determination error is larger, an effective solution is not provided at present.
Disclosure of Invention
The application mainly aims to provide a method and a device for determining an orbit of a spacecraft, a processor and electronic equipment, so as to solve the problem that acceleration generated by unloading of a momentum wheel of the spacecraft in the related art is mainly obtained by calculating a uniform acceleration model under an orbit coordinate system or a celestial sphere reference system, so that unloading acceleration is not stably solved, and further an orbit determination error is large.
In order to achieve the above object, according to one aspect of the present application, there is provided a method of determining a spacecraft orbit. The method comprises the following steps: constructing a rotation matrix according to a target attitude quaternion corresponding to the target attitude of the target spacecraft, wherein the rotation matrix is used for mutual conversion between an inertial coordinate system and a body coordinate system; constructing a first partial derivative model of the unloading acceleration under the inertial coordinate system to the acceleration under the body coordinate system according to the rotation matrix, and selecting a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic; according to the second partial derivative model and the observation data of the target spacecraft, a state equation of the target spacecraft is constructed, and the state equation is subjected to integral solution to obtain the target acceleration; and determining the orbit of the target spacecraft according to the target acceleration.
Further, before constructing a rotation matrix of the inertial coordinate system and the body coordinate system according to the target attitude quaternion of the target attitude of the target spacecraft, the method further comprises: and when the attitude of the target spacecraft is detected to be adjusted to the target attitude, acquiring a target attitude quaternion corresponding to the target attitude of the target spacecraft.
Further, before obtaining a target attitude quaternion corresponding to the target attitude of the target spacecraft, the method further includes: detecting whether the target spacecraft is in an active unloading mode; and if the target spacecraft is in an active unloading mode, sending a posture adjustment instruction to the target spacecraft so as to adjust the posture of the target spacecraft to the target posture.
Further, performing integral solution on the state equation to obtain the target acceleration includes: dividing the state equation into a state equation in an unloading state and a state equation in a non-unloading state; and carrying out integral solution on the state equation in the unloading state by a single-step integral method, and carrying out integral solution on the state equation in the non-unloading state by a multi-step integral method to obtain the target acceleration.
Further, the rotation matrix is:
Further, the first partial derivative model is:
In order to achieve the above object, according to another aspect of the present application, there is provided a determination apparatus of a spacecraft orbit. The device includes: the system comprises a first construction unit, a second construction unit and a third construction unit, wherein the first construction unit is used for constructing a rotation matrix according to a target attitude quaternion corresponding to a target attitude of a target spacecraft, and the rotation matrix is used for mutual conversion between an inertial coordinate system and a body coordinate system; the second construction unit is used for constructing a first partial derivative model of the unloading acceleration under the inertial coordinate system and the acceleration under the body coordinate system according to the rotation matrix, and selecting a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic; the third construction unit is used for constructing a state equation of the target spacecraft according to the second partial derivative model and the observation data of the target spacecraft, and performing integral solution on the state equation to obtain the target acceleration; and the determining unit determines the orbit of the target spacecraft according to the target acceleration.
Further, the apparatus further comprises: the acquiring unit is used for acquiring a target attitude quaternion corresponding to the target attitude of the target spacecraft when the attitude of the target spacecraft is detected to be adjusted to the target attitude before a rotation matrix of an inertia coordinate system and a body coordinate system is constructed according to the target attitude quaternion of the target attitude of the target spacecraft.
Further, the apparatus further comprises: the detection unit is used for detecting whether the target spacecraft is in an active unloading mode or not before acquiring a target attitude quaternion corresponding to a target attitude of the target spacecraft; and the sending unit is used for sending an attitude adjusting instruction to the target spacecraft to adjust the attitude of the target spacecraft to the target attitude if the target spacecraft is in the active unloading mode.
Further, the third building element comprises: the dividing module is used for dividing the state equation into a state equation in an unloading state and a state equation in a non-unloading state; and the solving module is used for carrying out integral solution on the state equation in the unloading state by a single-step integral method and carrying out integral solution on the state equation in the non-unloading state by a multi-step integral method to obtain the target acceleration.
Further, the rotation matrix is:
Further, the first partial derivative model is:
In order to achieve the above object, according to one aspect of the present application, there is provided a processor for executing a program, wherein the program is executed to perform the method for determining a spacecraft orbit described in any one of the above.
To achieve the above object, according to one aspect of the present application, there is provided an electronic device comprising one or more processors and a memory for storing one or more processors implementing the method for determining a spacecraft orbit of any one of the above.
Through the application, the following steps are adopted: constructing a rotation matrix according to a target attitude quaternion corresponding to a target attitude of the target spacecraft, wherein the rotation matrix is used for mutual conversion between an inertial coordinate system and a body coordinate system; constructing a first partial derivative model of the unloading acceleration under the inertial coordinate system to the acceleration under the body coordinate system according to the rotation matrix, and selecting a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic; according to the second partial derivative model and observation data of the target spacecraft, a state equation of the target spacecraft is constructed, and integral solution is carried out on the state equation to obtain target acceleration; the orbit of the target spacecraft is determined according to the target acceleration, and the problem that in the related technology, the acceleration generated by unloading of the spacecraft momentum wheel is mainly calculated by adopting an orbit coordinate system or a uniform acceleration model under a celestial sphere reference system, so that the unloading acceleration is not stably solved, and further the orbit determination error is large is solved. Establishing a rotation matrix between an inertial coordinate system and a body coordinate system through a target attitude quaternion and an angle to be rotated, establishing a partial derivative model for solving target acceleration for the body coordinate system based on the rotation matrix, establishing a state equation through the partial derivative model and observation data, and performing integral solution on the state equation to obtain target acceleration; and finally, determining the orbit of the target spacecraft according to the target acceleration, thereby achieving the effects of improving the solving accuracy of the unloading acceleration and reducing the error of orbit determination.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this application, are included to provide a further understanding of the application, and the description of the exemplary embodiments of the application are intended to be illustrative of the application and are not intended to limit the application. In the drawings:
fig. 1 is a flow chart of a method for determining a spacecraft orbit provided according to an embodiment of the present application;
FIG. 2 is a residual error map at the time of spacecraft offloading provided in accordance with an embodiment of the application;
FIG. 3 is a schematic illustration of position and velocity errors of a spacecraft provided in accordance with an embodiment of the present application;
FIG. 4 is a schematic diagram of orbit prediction for a spacecraft to orbit accuracy before next unloading according to an embodiment of the application;
fig. 5 is a schematic diagram of a spacecraft orbit determination apparatus provided in accordance with an embodiment of the present application;
fig. 6 is a schematic diagram of an electronic device provided according to an embodiment of the present application.
Detailed Description
It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict. The present application will be described in detail below with reference to the embodiments with reference to the attached drawings.
In order to make the technical solutions of the present application better understood by those skilled in the art, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application, and it is obvious that the described embodiments are only some embodiments of the present application, and not all embodiments. All other embodiments obtained by a person of ordinary skill in the art based on the embodiments in the present application without making any creative effort shall fall within the protection scope of the present application.
It should be noted that the terms "first," "second," and the like in the description and claims of this application and in the accompanying drawings are used for distinguishing between similar elements and not necessarily for describing a particular sequential or chronological order. It should be understood that the data so used may be interchanged under appropriate circumstances such that embodiments of the application described herein may be used. Furthermore, the terms "comprises," "comprising," and "having," and any variations thereof, are intended to cover a non-exclusive inclusion, such that a process, method, system, article, or apparatus that comprises a list of steps or elements is not necessarily limited to those steps or elements expressly listed, but may include other steps or elements not expressly listed or inherent to such process, method, article, or apparatus.
The present invention is described below with reference to preferred implementation steps, and fig. 1 is a flowchart of a method for determining a spacecraft orbit according to an embodiment of the present application, as shown in fig. 1, the method includes the following steps:
and S101, constructing a rotation matrix according to a target attitude quaternion corresponding to the target attitude of the target spacecraft, wherein the rotation matrix is used for mutual conversion between an inertial coordinate system and a body coordinate system.
Step S102, a first partial derivative model of unloading acceleration under an inertial coordinate system and acceleration under a body coordinate system is constructed according to the rotation matrix, and a second partial derivative model in the target direction is selected from the first partial derivative model according to the target control characteristics.
And S103, constructing a state equation of the target spacecraft according to the second partial derivative model and the observation data of the target spacecraft, and performing integral solution on the state equation to obtain the target acceleration.
And step S104, determining the orbit of the target spacecraft according to the target acceleration.
Since the motion of the spacecraft is usually described in an inertial coordinate system, the modeling of the acceleration generated by unloading mainly adopts an equivalent uniform acceleration model, i.e. it is assumed that the acceleration generated by unloading is constant in three directions of the inertial coordinate system in one unloading period. However, most engines used for unloading are attitude control engines, and thrust generated by unloading of most spacecrafts is usually concentrated in a certain direction of a body coordinate system according to the design of the installation position and the working mode of the attitude control engines.
In the technical background, a method for determining the orbit of the spacecraft in the active unloading mode is provided, unloading components in three directions in an inertial coordinate system are converted to a body coordinate system by using a conversion matrix, and the orbit of the spacecraft is determined by solving the acceleration in one direction of the body coordinate system.
Specifically, a rotation matrix between an inertial coordinate system and a body coordinate system is established through a target attitude quaternion and a to-be-rotated angle, a partial derivative model for solving target acceleration for the body coordinate system is established based on the rotation matrix, a state equation is established through the partial derivative model and observation data, and the state equation is subjected to integral solution to obtain target acceleration; and determining the orbit of the target spacecraft according to the target acceleration.
1. Establishing a rotation matrix M from an inertial coordinate system to a body coordinate system through an attitude quaternion
In the active unloading mode, the attitude of the target spacecraft is adjusted to a target attitude before unloading, and if a target attitude quaternion is Q, an inertial coordinate system is established to a rotation matrix M of a body coordinate system through the target attitude quaternion.
According to quaternion Q = [ Q ] 0 ，q 1 ，q 2 ，q 3 ]An expression of the rotation matrix M is obtained as follows:
wherein r is the angle to be rotated of the target spacecraft and is contained in four elements Q of the attitude, I is a unit matrix,
the rotation matrix is obtained by calculating equation (1):
2. establishing a partial derivative model of unloading acceleration under an inertial coordinate system to acceleration under a body coordinate system
Under the condition of the inertia space, the air conditioner,
wherein the content of the first and second substances,
represents the acceleration under the inertial coordinate system,
representing the acceleration, M, in a body coordinate system -1 Representing a rotation matrix, and obtaining the first partial derivative model by calculating the partial derivative of the acceleration in the body coordinate system as follows:
the thrust generated by the attitude control engine during active unloading is mainly concentrated in one direction of the body coordinate system, so that only the acceleration in one direction of the body coordinate system (corresponding to the second partial derivative model for selecting the target direction from the first partial derivative models according to the target control characteristics) can be considered during acceleration solving, for example, the target control characteristics can be set to select the partial derivative model in the z direction.
If the thrust is in the x direction of the body coordinate system, the second partial derivative model is:
if the thrust is in the y direction of the body coordinate system, the second partial derivative model is:
if the thrust is in the z direction of the body coordinate system, the second partial derivative model is:
3. orbit determination for a target spacecraft
Establishing a state equation through the second partial derivative model and observation data of the target spacecraft, and performing integral solution on the state equation to obtain target acceleration; and finally, determining the orbit of the target spacecraft according to the target acceleration. The observation data of the target spacecraft are generally the observation data of three days before and after the active unloading of the target spacecraft.
In conclusion, the second partial derivative model in the body coordinate system is established, and the target acceleration is solved by using the second partial derivative model, so that the accuracy of solving the acceleration is improved, and the accuracy of determining the spacecraft orbit is further improved.
In order to improve the accuracy of the spacecraft orbit maintenance through the unloading force, before constructing the rotation matrix of the inertial coordinate system and the body coordinate system according to the target attitude quaternion of the target attitude of the target spacecraft, the following processing is required: and when the attitude of the target spacecraft is detected to be adjusted to the target attitude, acquiring a target attitude quaternion corresponding to the target attitude of the target spacecraft. Detecting whether the target spacecraft is in an active unloading mode; and if the target spacecraft is in the active unloading mode, sending a posture adjustment instruction to the target spacecraft so as to adjust the posture of the target spacecraft to the target posture.
Specifically, generally, when the spacecraft is unloaded, the attitude of the target spacecraft is required to be in the target attitude, so that whether the target spacecraft is in the target attitude or not is detected firstly, and if the target spacecraft is in the target attitude, a target attitude quaternion corresponding to the target attitude is obtained. If the target spacecraft is not in the target attitude, the attitude of the target spacecraft needs to be adjusted. Firstly, detecting whether a target spacecraft is in an active unloading mode, and sending a posture adjusting instruction to the target spacecraft when the target spacecraft is in the active unloading mode; and then the target spacecraft adjusts the attitude to the target attitude according to the attitude adjustment instruction.
When the orbit maintenance of the spacecraft is required through the unloading force, the unloading force at this time is required to clearly generate a large acting force on the spacecraft, or the spacecraft is required to rotate under the action of the unloading force at this time.
And establishing a conversion matrix between an inertial coordinate system and a body coordinate system through the obtained target attitude quaternion. The attitude of the spacecraft is adjusted to the target attitude, so that the accuracy of calculating the acceleration is improved, and the accuracy of determining the orbit is further improved.
In order to improve the accuracy of integral solution of a state equation, in the method for determining a spacecraft orbit provided by the embodiment of the application, the state equation is subjected to integral solution in the following manner to obtain a target acceleration: dividing the state equation into a state equation in an unloading state and a state equation in a non-unloading state; and carrying out integral solution on the state equation in the unloading state by a single-step integral method, and carrying out integral solution on the state equation in the non-unloading state by a multi-step integral method to obtain the target acceleration.
Specifically, since the state of the target spacecraft changes relatively quickly in the unloaded state, and the state of the target spacecraft changes relatively slowly in the unloaded state, if the same integral solving method is adopted, the accuracy of solving acceleration is relatively low. According to the spacecraft orbit determination method, the state equation is divided into the state equation in the unloading state and the state equation in the non-unloading state, the state equation in the unloading state is solved by adopting a single-step integration method, the state equation in the non-unloading state is solved by adopting a multi-step integration method, and the resolving precision and accuracy of unloading acceleration are improved.
Optionally, in the method for determining a spacecraft orbit provided in the embodiment of the present application, the rotation matrix is:
Optionally, in the method for determining a spacecraft orbit provided in the embodiment of the present application, the first partial derivative model is:
wherein, the first and the second end of the pipe are connected with each other,
represents the acceleration in the inertial coordinate system,
the acceleration in the body coordinate system is shown, and x, y and z are coordinate axes in different directions.
In an alternative embodiment, the determination of the precision orbit is performed by using 2 ways of solving 3 direction empirical forces of the inertial coordinate system before improvement and solving only z direction empirical force of the body coordinate system after improvement (i.e. the orbit determination method provided by the application). The Root Mean Square (RMS) and form error are counted in two ways in table 1. As shown in table 1, the RMS values calculated by the 2 methods are equivalent in terms of calculation effect, but the improved form error is obviously better than that before the improvement because the improved calculation parameters are more reasonable.
TABLE 1
In an alternative embodiment, fig. 2 shows a residual map of 6 unloads, which are solved by the improved method, and the residual map corresponds to sub-maps (a) - (f) in sequence in time. The accuracy of the overlapped arc segments is evaluated by using an overlapped arc segment comparison method, overlapped parts of the adjacent 2 precise tracks are compared, position and speed errors are counted, and the result is shown in fig. 3. As can be seen from FIG. 3, the position accuracy of the improved overlapped arc segment is less than 1km, the speed accuracy is better than 4mm/s, and is obviously better than the accuracy before the improvement.
In an alternative embodiment, fig. 4 shows the track accuracy before the track is predicted to be unloaded next time by using the data before unloading and the data 1.5 days after unloading, the accuracy is obviously improved by using the improved method, the predicted position accuracy is 1.5km, and the speed accuracy is better than 1cm/s.
According to the method for determining the spacecraft orbit, a rotation matrix is constructed according to a target attitude quaternion corresponding to a target attitude of a target spacecraft, wherein the rotation matrix is used for mutual conversion between an inertial coordinate system and a body coordinate system; constructing a first partial derivative model of the unloading acceleration under the inertial coordinate system to the acceleration under the body coordinate system according to the rotation matrix, and selecting a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic; according to the second partial derivative model and observation data of the target spacecraft, a state equation of the target spacecraft is constructed, and integral solution is carried out on the state equation to obtain target acceleration; the orbit of the target spacecraft is determined according to the target acceleration, and the problem that the unloading acceleration is not stably solved due to the fact that the acceleration generated by unloading of the spacecraft momentum wheel in the related technology is mainly calculated by adopting a uniform acceleration model under an orbit coordinate system or a celestial sphere reference system is solved. Establishing a rotation matrix between an inertial coordinate system and a body coordinate system through a target attitude quaternion and an angle to be rotated, establishing a partial derivative model for solving target acceleration for the body coordinate system based on the rotation matrix, establishing a state equation through the partial derivative model and observation data, and performing integral solution on the state equation to obtain target acceleration; and determining the orbit of the target spacecraft according to the target acceleration, thereby achieving the effects of improving the solving accuracy of the unloading acceleration and reducing the error of orbit determination.
It should be noted that the steps illustrated in the flowcharts of the figures may be performed in a computer system such as a set of computer-executable instructions and that, although a logical order is illustrated in the flowcharts, in some cases, the steps illustrated or described may be performed in an order different than presented herein.
The embodiment of the present application further provides a device for determining a spacecraft orbit, and it should be noted that the device for determining a spacecraft orbit of the embodiment of the present application can be used to execute the method for determining a spacecraft orbit provided by the embodiment of the present application. The following describes a spacecraft orbit determination device provided in an embodiment of the present application.
Fig. 5 is a schematic diagram of a spacecraft orbit determination apparatus according to an embodiment of the application. As shown in fig. 5, the apparatus includes: a first building element 501, a second building element 502, a third building element 503 and a determination element 504.
The first constructing unit 501 is configured to construct a rotation matrix according to a target attitude quaternion corresponding to a target attitude of the target spacecraft, where the rotation matrix is used for interconversion between an inertial coordinate system and a body coordinate system.
The second constructing unit 502 is configured to construct a first partial derivative model of the unloading acceleration in the inertial coordinate system with respect to the acceleration in the body coordinate system according to the rotation matrix, and select a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic.
The third constructing unit 503 is configured to construct a state equation of the target spacecraft according to the second partial derivative model and the observation data of the target spacecraft, and perform integral solution on the state equation to obtain a target acceleration.
A determining unit 504, configured to determine an orbit of the target spacecraft according to the target acceleration.
The spacecraft orbit determination apparatus provided by the embodiment of the application is configured to construct a rotation matrix according to a target attitude quaternion corresponding to a target attitude of a target spacecraft through the first construction unit 501, wherein the rotation matrix is used for interconversion between an inertial coordinate system and a body coordinate system. The second constructing unit 502 is configured to construct a first partial derivative model of the unloading acceleration in the inertial coordinate system with respect to the acceleration in the body coordinate system according to the rotation matrix, and select a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic. The third constructing unit 503 is configured to construct a state equation of the target spacecraft according to the second partial derivative model and the observation data of the target spacecraft, and perform integral solution on the state equation to obtain a target acceleration. The determining unit 504 is configured to determine the orbit of the target spacecraft according to the target acceleration, and solve the problem in the related art that the acceleration generated by unloading the spacecraft momentum wheel is mainly calculated by using a uniform acceleration model in an orbit coordinate system or a celestial sphere reference system, which causes unstable unloading acceleration solution and further causes a large orbit determination error. Establishing a rotation matrix between an inertial coordinate system and a body coordinate system through a target attitude quaternion and an angle to be rotated, establishing a partial derivative model for solving target acceleration for the body coordinate system based on the rotation matrix, establishing a state equation through the partial derivative model and observation data, and performing integral solution on the state equation to obtain target acceleration; and determining the orbit of the target spacecraft according to the target acceleration, thereby achieving the effects of improving the accuracy of solving the unloading acceleration and reducing the error of orbit determination.
Optionally, in the apparatus for determining a spacecraft orbit provided in the embodiment of the present application, the apparatus further includes: the acquiring unit is used for acquiring a target attitude quaternion corresponding to the target attitude of the target spacecraft when the attitude of the target spacecraft is detected to be adjusted to the target attitude before the rotation matrix of the inertial coordinate system and the body coordinate system is constructed according to the target attitude quaternion of the target attitude of the target spacecraft.
Optionally, in the apparatus for determining a spacecraft orbit provided in the embodiment of the present application, the apparatus further includes: the detection unit is used for detecting whether the target spacecraft is in an active unloading mode or not before acquiring a target attitude quaternion corresponding to the target attitude of the target spacecraft; and the sending unit is used for sending a posture adjusting instruction to the target spacecraft to adjust the posture of the target spacecraft to the target posture if the target spacecraft is in the active unloading mode.
Optionally, in the apparatus for determining a spacecraft orbit provided in the embodiment of the present application, the third building unit 503 includes: the dividing module is used for dividing the state equation into a state equation in an unloading state and a state equation in a non-unloading state; and the solving module is used for carrying out integral solution on the state equation in the unloading state by a multi-component integration method and carrying out integral solution on the state equation in the non-unloading state by a single-step integration method to obtain the target acceleration.
Optionally, in the apparatus for determining a spacecraft orbit provided in the embodiment of the present application, the rotation matrix is:
where M is a rotation matrix, Q = [ Q ] 0 ，q 1 ，q 2 ，q 3 ]For the target attitude quaternion, q0, q1, q2, q3 are four attitude parameters consisting of euler axis/angle parameters, respectively.
Optionally, in the apparatus for determining a spacecraft orbit provided in the embodiment of the present application, the first partial derivative model is:
wherein the content of the first and second substances,
represents the acceleration in the inertial coordinate system,
the acceleration in the body coordinate system is shown, and x, y and z are coordinate axes in different directions.
The spacecraft orbit determination device comprises a processor and a memory, wherein the first building unit 501, the second building unit 502, the third building unit 503, the determination unit 504 and the like are stored in the memory as program units, and the corresponding functions are realized by executing the program units stored in the memory by the processor.
The processor comprises a kernel, and the kernel calls the corresponding program unit from the memory. The kernel can be set to be one or more, and the orbit determination of the spacecraft is realized by adjusting the kernel parameters.
The memory may include volatile memory in a computer readable medium, random Access Memory (RAM) and/or nonvolatile memory such as Read Only Memory (ROM) or flash memory (flash RAM), and the memory includes at least one memory chip.
The embodiment of the invention provides a processor, which is used for running a program, wherein the program executes a spacecraft orbit determination method during running.
As shown in fig. 6, an embodiment of the present invention provides an electronic device, where the device includes a processor, a memory, and a program stored in the memory and executable on the processor, and the processor executes the program to implement the following steps: constructing a rotation matrix according to a target attitude quaternion corresponding to the target attitude of the target spacecraft, wherein the rotation matrix is used for mutual conversion between an inertial coordinate system and a body coordinate system; constructing a first partial derivative model of the unloading acceleration under the inertial coordinate system to the acceleration under the body coordinate system according to the rotation matrix, and selecting a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic; according to the second partial derivative model and observation data of the target spacecraft, a state equation of the target spacecraft is constructed, and integral solution is carried out on the state equation to obtain target acceleration; and determining the orbit of the target spacecraft according to the target acceleration.
Optionally, before constructing the rotation matrix of the inertial coordinate system and the body coordinate system according to the target attitude quaternion of the target attitude of the target spacecraft, the method further includes: and when the attitude of the target spacecraft is detected to be adjusted to the target attitude, acquiring a target attitude quaternion corresponding to the target attitude of the target spacecraft.
Optionally, before obtaining a target attitude quaternion corresponding to the target attitude of the target spacecraft, the method further includes: detecting whether the target spacecraft is in an active unloading mode; and if the target spacecraft is in the active unloading mode, sending a posture adjustment instruction to the target spacecraft so as to adjust the posture of the target spacecraft to the target posture.
Optionally, the integral solution of the state equation to obtain the target acceleration includes: dividing the state equation into a state equation in an unloading state and a state equation in a non-unloading state; and performing integral solution on the state equation in the unloading state by a single-step integral method, and performing integral solution on the state equation in the non-unloading state by a multi-step integral method to obtain the target acceleration.
Optionally, the rotation matrix is:
Optionally, the first partial derivative model is:
The device herein may be a server, a PC, a PAD, a mobile phone, etc.
The present application further provides a computer program product adapted to perform a program for initializing the following method steps when executed on a data processing device: constructing a rotation matrix according to a target attitude quaternion corresponding to a target attitude of the target spacecraft, wherein the rotation matrix is used for mutual conversion between an inertial coordinate system and a body coordinate system; constructing a first partial derivative model of the unloading acceleration under the inertial coordinate system to the acceleration under the body coordinate system according to the rotation matrix, and selecting a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic; according to the second partial derivative model and observation data of the target spacecraft, a state equation of the target spacecraft is constructed, and integral solution is carried out on the state equation to obtain target acceleration; and determining the orbit of the target spacecraft according to the target acceleration.
Optionally, before constructing the rotation matrix of the inertial coordinate system and the body coordinate system according to the target attitude quaternion of the target attitude of the target spacecraft, the method further includes: and when the attitude of the target spacecraft is detected to be adjusted to the target attitude, acquiring a target attitude quaternion corresponding to the target attitude of the target spacecraft.
Optionally, before obtaining a target attitude quaternion corresponding to the target attitude of the target spacecraft, the method further includes: detecting whether the target spacecraft is in an active unloading mode; and if the target spacecraft is in the active unloading mode, sending a posture adjustment instruction to the target spacecraft so as to adjust the posture of the target spacecraft to the target posture.
Optionally, the integral solution of the state equation to obtain the target acceleration includes: dividing the state equation into a state equation in an unloading state and a state equation in a non-unloading state; and performing integral solution on the state equation in the unloading state by a single-step integral method, and performing integral solution on the state equation in the non-unloading state by a multi-step integral method to obtain the target acceleration.
Optionally, the rotation matrix is:
Optionally, the first partial derivative model is:
As will be appreciated by one skilled in the art, embodiments of the present application may be provided as a method, system, or computer program product. Accordingly, the present application may take the form of an entirely hardware embodiment, an entirely software embodiment or an embodiment combining software and hardware aspects. Furthermore, the present application may take the form of a computer program product embodied on one or more computer-usable storage media (including, but not limited to, disk storage, CD-ROM, optical storage, and so forth) having computer-usable program code embodied therein.
The present application is described with reference to flowchart illustrations and/or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the application. It will be understood that each flow and/or block of the flow diagrams and/or block diagrams, and combinations of flows and/or blocks in the flow diagrams and/or block diagrams, can be implemented by computer program instructions. These computer program instructions may be provided to a processor of a general purpose computer, special purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, create means for implementing the functions specified in the flowchart flow or flows and/or block diagram block or blocks.
These computer program instructions may also be stored in a computer-readable memory that can direct a computer or other programmable data processing apparatus to function in a particular manner, such that the instructions stored in the computer-readable memory produce an article of manufacture including instruction means which implement the function specified in the flowchart flow or flows and/or block diagram block or blocks.
These computer program instructions may also be loaded onto a computer or other programmable data processing apparatus to cause a series of operational steps to be performed on the computer or other programmable apparatus to produce a computer implemented process such that the instructions which execute on the computer or other programmable apparatus provide steps for implementing the functions specified in the flowchart flow or flows and/or block diagram block or blocks.
In a typical configuration, a computing device includes one or more processors (CPUs), input/output interfaces, network interfaces, and memory.
The memory may include forms of volatile memory in a computer readable medium, random Access Memory (RAM) and/or non-volatile memory, such as Read Only Memory (ROM) or flash memory (flash RAM). The memory is an example of a computer-readable medium.
Computer-readable media, including both non-transitory and non-transitory, removable and non-removable media, may implement information storage by any method or technology. The information may be computer readable instructions, data structures, modules of a program, or other data. Examples of computer storage media include, but are not limited to, phase change memory (PRAM), static Random Access Memory (SRAM), dynamic Random Access Memory (DRAM), other types of Random Access Memory (RAM), read Only Memory (ROM), electrically Erasable Programmable Read Only Memory (EEPROM), flash memory or other memory technology, compact disc read only memory (CD-ROM), digital Versatile Discs (DVD) or other optical storage, magnetic cassettes, magnetic tape, magnetic disk storage or other magnetic storage devices, or any other non-transmission medium that can be used to store information that can be accessed by a computing device. As defined herein, a computer readable medium does not include a transitory computer readable medium such as a modulated data signal and a carrier wave.
It should also be noted that the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus. Without further limitation, an element defined by the phrases "comprising one of 8230; \8230;" 8230; "does not exclude the presence of additional like elements in a process, method, article, or apparatus that comprises the element.
As will be appreciated by one skilled in the art, embodiments of the present application may be provided as a method, system, or computer program product. Accordingly, the present application may take the form of an entirely hardware embodiment, an entirely software embodiment or an embodiment combining software and hardware aspects. Furthermore, the present application may take the form of a computer program product embodied on one or more computer-usable storage media (including, but not limited to, disk storage, CD-ROM, optical storage, and the like) having computer-usable program code embodied therein.
The above are merely examples of the present application and are not intended to limit the present application. Various modifications and changes may occur to those skilled in the art. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present application should be included in the scope of the claims of the present application.
Claims (9)
1. A method for determining a spacecraft orbit, comprising:
constructing a rotation matrix according to a target attitude quaternion corresponding to a target attitude of the target spacecraft, wherein the rotation matrix is used for mutual conversion between an inertial coordinate system and a body coordinate system;
constructing a first partial derivative model of unloading acceleration under an inertial coordinate system to acceleration under a body coordinate system according to the rotation matrix, and selecting a second partial derivative model in a target direction from the first partial derivative model according to target control characteristics;
according to the second partial derivative model and the observation data of the target spacecraft, a state equation of the target spacecraft is constructed, and integral solution is carried out on the state equation to obtain target acceleration;
and determining the orbit of the target spacecraft according to the target acceleration.
2. The method of claim 1, wherein prior to constructing the rotation matrix of the inertial and body coordinate systems from the target attitude quaternion for the target attitude of the target spacecraft, the method further comprises:
and when the attitude of the target spacecraft is detected to be adjusted to the target attitude, acquiring a target attitude quaternion corresponding to the target attitude of the target spacecraft.
3. The method of claim 2, wherein prior to obtaining a target attitude quaternion for the target attitude of the target spacecraft, the method further comprises:
detecting whether the target spacecraft is in an active unloading mode;
and if the target spacecraft is in an active unloading mode, sending an attitude adjusting instruction to the target spacecraft so as to adjust the attitude of the target spacecraft to the target attitude.
4. The method of claim 1, wherein solving the state equation by integration to obtain the target acceleration comprises:
dividing the state equation into a state equation in an unloading state and a state equation in a non-unloading state;
and carrying out integral solution on the state equation in the unloading state by a single-step integral method, and carrying out integral solution on the state equation in the non-unloading state by a multi-step integral method to obtain the target acceleration.
5. The method of claim 1, wherein the rotation matrix is:
wherein M is the rotation matrix, Q = [ Q ] 0 ，q 1 ，q 2 ，q 3 ]And q0, q1, q2 and q3 are four attitude parameters consisting of Euler axis/angle parameters respectively, which are the target attitude quaternion.
6. The method of claim 1, wherein the first partial derivative model is:
wherein, the first and the second end of the pipe are connected with each other,
represents the acceleration in the inertial coordinate system,
and the acceleration under the body coordinate system is represented, and x, y and z are coordinate axes under different directions.
7. An apparatus for determining a spacecraft orbit, comprising:
the system comprises a first construction unit, a second construction unit and a third construction unit, wherein the first construction unit is used for constructing a rotation matrix according to a target attitude quaternion corresponding to a target attitude of a target spacecraft, and the rotation matrix is used for mutual conversion between an inertial coordinate system and a body coordinate system;
the second construction unit is used for constructing a first partial derivative model of the unloading acceleration under the inertial coordinate system and the acceleration under the body coordinate system according to the rotation matrix, and selecting a second partial derivative model in the target direction from the first partial derivative model according to the target control characteristic;
the third construction unit is used for constructing a state equation of the target spacecraft according to the second partial derivative model and the observation data of the target spacecraft, and performing integral solution on the state equation to obtain target acceleration;
and the determining unit is used for determining the orbit of the target spacecraft according to the target acceleration.
8. A processor, characterized in that the processor is configured to run a program, wherein the program when running performs the method of determining a spacecraft orbit according to any one of claims 1 to 6.
9. An electronic device comprising one or more processors and memory for storing one or more programs, wherein the one or more programs, when executed by the one or more processors, cause the one or more processors to implement the method of determining a spacecraft orbit of any of claims 1-6.
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| CN116975501A (en) * | 2023-09-20 | 2023-10-31 | 中科星图测控技术股份有限公司 | Method for optimizing satellite load to ground target coverage calculation |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN116975501A (en) * | 2023-09-20 | 2023-10-31 | 中科星图测控技术股份有限公司 | Method for optimizing satellite load to ground target coverage calculation |
| CN116975501B (en) * | 2023-09-20 | 2023-12-15 | 中科星图测控技术股份有限公司 | Method for optimizing satellite load to ground target coverage calculation |
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