CN114415716B - Method, device and medium for maintaining constellation configuration - Google Patents
Method, device and medium for maintaining constellation configuration Download PDFInfo
- Publication number
- CN114415716B CN114415716B CN202111555151.XA CN202111555151A CN114415716B CN 114415716 B CN114415716 B CN 114415716B CN 202111555151 A CN202111555151 A CN 202111555151A CN 114415716 B CN114415716 B CN 114415716B
- Authority
- CN
- China
- Prior art keywords
- target point
- set target
- track
- orbit
- satellite
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 238000000034 method Methods 0.000 title claims abstract description 40
- 238000004364 calculation method Methods 0.000 claims abstract description 18
- 238000006243 chemical reaction Methods 0.000 claims description 7
- 238000004891 communication Methods 0.000 claims description 7
- 238000004590 computer program Methods 0.000 claims description 6
- 230000006870 function Effects 0.000 claims description 5
- 230000000630 rising effect Effects 0.000 claims description 5
- 230000001133 acceleration Effects 0.000 claims description 4
- 230000005484 gravity Effects 0.000 claims description 4
- 239000000126 substance Substances 0.000 claims description 4
- 230000009466 transformation Effects 0.000 claims description 4
- 150000001875 compounds Chemical class 0.000 claims description 2
- 230000008569 process Effects 0.000 claims description 2
- 238000010304 firing Methods 0.000 claims 3
- 239000000243 solution Substances 0.000 description 14
- 238000012545 processing Methods 0.000 description 7
- 238000012423 maintenance Methods 0.000 description 6
- 230000008859 change Effects 0.000 description 4
- 238000010586 diagram Methods 0.000 description 4
- 230000001360 synchronised effect Effects 0.000 description 4
- 238000009825 accumulation Methods 0.000 description 3
- 238000002347 injection Methods 0.000 description 3
- 239000007924 injection Substances 0.000 description 3
- 230000005540 biological transmission Effects 0.000 description 2
- 238000005259 measurement Methods 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 230000009471 action Effects 0.000 description 1
- 238000003491 array Methods 0.000 description 1
- 238000004422 calculation algorithm Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000014509 gene expression Effects 0.000 description 1
- 230000007774 longterm Effects 0.000 description 1
- 230000003287 optical effect Effects 0.000 description 1
Images
Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Position Fixing By Use Of Radio Waves (AREA)
Abstract
The embodiment of the invention discloses a method, a device and a medium for maintaining constellation configuration; the method comprises the following steps: receiving nominal orbit parameters and timestamps annotated by a ground station; when the satellite receives the mth time passing through the set target point, obtaining an average nominal orbit parameter in a time period from the time stamp of receiving the upper note to the mth time passing through the set target point through recursive calculation according to the nominal orbit parameter of the upper note; acquiring the corresponding orbit deviation amount when the mth time passes through the set target point according to the average nominal orbit parameter and the satellite orbit parameter when the mth time passes through the set target point; when the number m of times of passing through the set target point is greater than a set number threshold, selecting a latitude argument, a speed increment and an ignition duration of an ignition position according to the track deviation amount and the track control target type; and igniting according to the latitude argument, the speed increment and the ignition duration of the ignition position to complete the track control.
Description
Technical Field
The embodiment of the invention relates to the technical field of satellite operation control, in particular to a method, a device and a medium for maintaining constellation configuration.
Background
Low orbit satellites are generally circular orbits, which have attracted great interest to giant constellation designers and operators due to their characteristics of low orbital height, short transmission delay, low path loss, and the like.
The on-orbit satellite in the constellation is affected by perturbation power such as earth oblateness perturbation, sun-moon attraction perturbation, atmospheric resistance perturbation, sunlight pressure and the like besides the action of earth gravity; moreover, due to different in-orbit deviation and different in perturbation power of each satellite in the constellation, the satellites in the constellation are subjected to different degrees of orbital attenuation and phase drift among the satellites, so that the constellation configuration is diverged. If the constellation configuration is not maintained, the constellation structure will be out of order and the service performance of the constellation will be affected.
In order to maintain the constellation configuration, a conventional scheme generally selects a reference satellite through orbit data of each satellite in the constellation; then, calculating the speed increment of other satellites (such as the starting time and the starting place of the thruster) except the reference satellite in the constellation; and (4) injecting the startup time of the thruster to other satellites through the ground station to maintain the constellation configuration.
Based on the explanation of the conventional scheme, if the control quantity calculation and the upper injection are completely carried out by depending on the ground station, the difficulty of the ground constellation management is increased, and the constellation configuration is difficult to be accurately maintained due to the time delay caused by the upper injection time, so that the configuration maintaining effect of the satellite constellation is influenced.
Disclosure of Invention
In view of the above, embodiments of the present invention are directed to a method, an apparatus, and a medium for maintaining a constellation configuration; the control difficulty of the ground station on the large-scale constellation configuration can be reduced, and the control precision of the satellite constellation configuration maintenance is improved.
The technical scheme of the embodiment of the invention is realized as follows:
in a first aspect, an embodiment of the present invention provides a method for maintaining a constellation configuration, where the method is applied to any one of satellites in a constellation except a reference satellite, and the method includes:
receiving nominal orbit parameters and timestamps annotated by a ground station;
when the satellite receives the mth time passing through the set target point, obtaining an average nominal orbit parameter in a time period from the time stamp of receiving the upper note to the mth time passing through the set target point through recursive calculation according to the nominal orbit parameter of the upper note;
acquiring the corresponding orbit deviation amount when the mth time passes through the set target point according to the average nominal orbit parameter and the satellite orbit parameter when the mth time passes through the set target point;
when the number m of times of passing through the set target point is greater than a set number threshold, selecting a latitude argument, a speed increment and an ignition duration of an ignition position according to the track deviation amount and the track control target type;
and igniting according to the latitude argument, the speed increment and the ignition time length of the ignition position to complete the track control.
In a second aspect, an embodiment of the present invention provides an apparatus for maintaining a constellation configuration, where the apparatus is applied to any one of other satellites in the constellation except for a reference satellite, and the apparatus includes: a receiving part, a recursion calculating part, an obtaining part, a selecting part and an ignition part; wherein, the first and the second end of the pipe are connected with each other,
the receiving portion configured to receive nominal orbit parameters and timestamps annotated by a ground station;
the recursion calculation part is configured to obtain an average nominal orbit parameter in a period from the time stamp of the received upper note to the mth passing of the set target point through recursion calculation according to the nominal orbit parameter of the upper note when the satellite receives the mth passing of the set target point;
the acquisition part is configured to acquire the corresponding orbit deviation amount when the set target point is passed the mth time according to the average nominal orbit parameter and the orbit parameter of the satellite when the set target point is passed the mth time;
the selecting part is configured to select a latitude argument, a speed increment and an ignition duration of the ignition position according to the track deviation amount and the track control target type when the number m of times of passing through the set target point is greater than a set number threshold;
the ignition part is configured to ignite according to the latitude argument, the speed increment and the ignition time length of the ignition position so as to complete the track control.
In a third aspect, an embodiment of the present invention provides an on-board computing device, where the on-board computing device includes: a communication interface, a memory and a processor; the various components are coupled together by a bus system; wherein the content of the first and second substances,
the communication interface is used for receiving and sending signals in the process of receiving and sending information with other external network elements;
the memory for storing a computer program operable on the processor;
the processor is configured to, when executing the computer program, perform the steps of the method for maintaining a constellation configuration of the first aspect.
In a fourth aspect, an embodiment of the present invention provides a computer storage medium storing a program for maintaining a constellation configuration, where the program for maintaining a constellation configuration implements the steps of the method for maintaining a constellation configuration in the first aspect when executed by at least one processor.
The embodiment of the invention provides a method, a device and a medium for maintaining constellation configuration; the ground station only needs to inject the nominal orbit parameters and the time stamps of the satellites, then the satellites complete the recursion of the nominal orbit through the on-satellite system of the satellites, acquire the orbit deviation amount through multi-turn accumulation, and then complete the orbit control based on the orbit deviation amount. The configuration can be autonomously maintained on the satellite without complex calculation processing, the control difficulty of the ground station on a large-scale constellation configuration is greatly reduced, and the control precision of the maintenance of the satellite constellation configuration is improved.
Drawings
Fig. 1 is a schematic flowchart of a method for maintaining a constellation configuration according to an embodiment of the present invention;
FIG. 2 is a schematic diagram of an RSW coordinate system provided by an embodiment of the present invention;
fig. 3 is a schematic diagram illustrating an apparatus for maintaining a constellation configuration according to an embodiment of the present invention;
fig. 4 is a schematic diagram of a hardware structure of a satellite-borne computing device according to an embodiment of the present invention.
Detailed Description
The technical solution in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention.
The embodiment of the invention is expected to adopt a multi-circle accumulation mode to calculate the self latitude argument change among the satellites in the constellation, and adopt an orbit recursion mode to compare the latitude argument change rate deviation and the latitude argument deviation to obtain the control data of each satellite, thereby realizing that the satellites finish respective orbit configuration maintenance on the satellite independently and further realizing the constellation configuration maintenance on the satellite.
Based on this, referring to fig. 1, a method for maintaining a constellation configuration provided by an embodiment of the present invention is shown, which may be applied to an on-board system of any one of other satellites in a constellation except a reference satellite, such as an on-board computer, and may include:
s101: receiving nominal orbit parameters and timestamps annotated by a ground station;
s102: when the satellite receives the mth time passing through the set target point, the average nominal orbit parameter in the time period from the time stamp of receiving the upper note to the mth time passing through the set target point is obtained through recursive calculation according to the nominal orbit parameter of the upper note;
s103: acquiring the corresponding orbit deviation amount when the mth time passes through the set target point according to the average nominal orbit parameter and the satellite orbit parameter when the mth time passes through the set target point;
s104: when the number m of times of passing through the set target point is greater than a set number threshold, selecting a latitude argument, a speed increment and an ignition duration of an ignition position according to the track deviation amount and the track control target type;
s105: and igniting according to the latitude argument, the speed increment and the ignition duration of the ignition position to complete the track control.
Through the technical scheme shown in fig. 1, the ground station only needs to inject the nominal orbit parameters and the time stamps of the satellites, then the satellites complete recursion of the nominal orbit through a new system of the ground station, acquire the orbit deviation amount through multi-turn accumulation, and then complete orbit control based on the orbit deviation amount. The configuration can be autonomously maintained on the satellite without complex calculation processing, the control difficulty of the ground station on large-scale constellation configuration is greatly reduced, and the control precision of satellite constellation configuration maintenance is improved.
For the technical solution shown in fig. 1, in some examples, the track parameter is preferably six classical tracks, which specifically includes: the method comprises the following steps of (1) determining a semi-major axis a of a track, an eccentricity e of the track, a track inclination angle i, a rising intersection declination omega, a perigee argument omega and a mean perigee angle M; therefore, by combining the injection timestamp, the information amount injected to the satellite by the ground station is small, the ground station does not need to perform complex calculation processing, and the control difficulty of the ground station on a large-scale constellation configuration is reduced. Second, in some examples, the set target point is preferably a descending point. The descent intersection point is an intersection point of the orbital plane and the equatorial plane when the satellite travels from north to south.
Based on the above example, it should be noted that, from the nominal orbit parameters and the time stamps of the satellite receiving the upper notes, the satellite counts each time it runs and passes the lower crossing point. While the satellite can obtain the instantaneous nominal orbit parameter when passing through the descending intersection point by the aid of the on-satellite computer of the satellite according to the recursion of the annotated nominal orbit parameter as the satellite operates, in the embodiment of the invention, the average nominal orbit parameter in the time period delta t from the time stamp of the received annotation to the ith passing of the set target point is preferably used for representingInstantaneous nominal track parameters at the i-th pass through the descent intersection point. In detail, consider J 2 The Kepler orbit perturbation analytic solution in perturbation is shown as formula 1:
in the above formula, if the parameters of the orbit semi-major axis a, the orbit eccentricity e, the orbit inclination angle i, the ascension point right ascension omega, the perigee argument omega and the mean perigee angle M are not provided with asterisks, the parameters are expressed as instantaneous orbit parameters; if the parameters of the track semi-major axis a, the track eccentricity e, the track inclination angle i, the ascent intersection declination omega, the perigee argument omega and the mean perigee angle M are provided with asterisks, the parameters are represented as track parameters comprising the change of long-term items; the delta-related parameter represents a short-period perturbation term of the order of O (J) 2 ). Based on the above formula, after the satellite receives the nominal instantaneous orbit parameter annotated by passing through the set target point for the ith time, the nominal instantaneous orbit parameter can be converted into a nominal average orbit parameter which can be used for orbit control quantity calculation through the on-board computer, specifically, short-period and long-period terms in the nominal instantaneous orbit parameter can be removed, so that the corresponding nominal average orbit parameter is obtained; it can be understood that, since the average orbit parameter is usually adopted to calculate the orbit control quantity, in the subsequent part of the embodiment of the present invention, after obtaining other instantaneous orbit parameters, the orbit control quantity is converted into corresponding orbit average parameters by using a set on-board algorithm through an on-board computer, which is not described in detail in the embodiment of the present invention.
In some possible implementations, the obtaining, by recursive computation, an average nominal orbit parameter in a period from the time when the upper note is received to the ith passing through the set target point according to the nominal orbit parameter of the upper note includes:
nominal average orbit parameter using nominal instantaneous orbit parameter conversion based on upper notes
And equation 2 is obtained from the reception of the timestamp of the upper note to the ith passage to the set targetAverage nominal orbit parameter over time period Δ t of a point
In the above-mentioned formula, the compound of formula,
and
respectively represent
The first differential of the function.
For the implementation mode, after the average nominal orbit parameter when the ith pass through the set target point is obtained, the orbit offset can be obtained according to the orbit parameter obtained by the satellite through the actual measurement of the sensor, and the average nominal orbit parameters of the satellite are set as a d 、e d 、i d 、Ω d 、ω d 、M d The orbit parameters obtained by the satellite through the actual measurement of a sensor are a, e, i, omega and M respectively; in some examples, the obtaining the orbit deviation amount corresponding to the mth time passing through the set target point according to the average nominal orbit parameter and the orbit parameter of the satellite at the mth time passing through the set target point includes:
acquiring the track deviation phase angle when the mth time passes through the set target point based on the formula 3
Wherein, M m Represents the fact that the target point is set by the mth passAverage mean anomaly, omega, of the mean anomaly transformation m Mean perigee argument, M, representing the transformation from the measured perigee argument at the mth pass through the set target point dm Denotes the mean nominal mean approximate point angle, ω, at the mth pass through the set target point dm Representing an average nominal perigee argument at the mth pass through the set target point;
according to the track deviation phase angle when the mth time passes through the set target point
And the number of times the satellite passes the set target point, and obtaining the orbit deviation phase angle speed when the mth time passes the set target point based on the formula 4
Wherein the content of the first and second substances,
represents the track deviation phase angle at the time of the 1 st pass through the set target point;
acquiring a track inclination angle deviation delta i and a track ascent intersection right ascension deviation delta omega when the mth time passes through the set target point according to equations 5 and 6:
Δi＝i m -i dm (5)
ΔΩ＝Ω m -Ω dm (6)
wherein i m Average track inclination angle i representing the conversion from the measured track inclination angle at the m-th pass through the set target point dm Denotes the mean nominal orbital inclination, Ω, at the m-th pass through the set target point m Mean rising point-right ascension, Ω, representing the conversion from the measured rising point-right ascension at the mth passage through the set target point dm Represents the average nominal ascent point ascent in the mth pass through the set target point.
It is understood that each time the satellite passes a descent intersection point, a corresponding orbit deviation amount can be obtained as shown in equations 3 to 6 in the above example. As the satellite operates, the satellite passes through the descending point for a plurality of times, and when the number of times of passing is accumulated to a certain number N (i.e., the set number threshold), it can be considered that the orbit deviation of the satellite is accumulated to a degree that cannot be ignored and needs to be adjusted. Based on this, in some possible implementations, the selecting the latitude argument, the speed increment, and the ignition duration of the ignition location according to the track deviation amount and the track control target type includes:
selecting a latitude argument u of the ignition position according to the track control target type and the formula 7:
according to the Gaussian perturbation equation of the on-orbit motion satellite in the RSW coordinate system and the orbit eccentricity e of the near-circular orbit being approximate to 0, a simplified perturbation equation shown in the formula 8 is obtained:
wherein, F S 、F W Perturbation thrust components along the transverse direction and the normal direction of the track under an RSW coordinate system are respectively; Δ v S 、Δv W Perturbation acceleration components along the transverse direction and the normal direction of the track under an RSW coordinate system are respectively; f is a true proximal angle;
is the track angular velocity; mu is an earth gravity constant; u = ω + f;
according to track deviation phase angle
Deviation from track phase angular velocity
And equation 9, acquiring the track angular velocity Δ n that needs to be changed:
wherein, T free Representing a drift time length;
based on the setting of the near-circular orbit, equation 10 is obtained:
Δn＝3Δv/a (10)
where Δ v represents the required velocity increment for the satellite;
combining equation 8 and equation 10, the velocity increment Δ v required to obtain a satellite is shown in equation 11:
calculating and obtaining the ignition duration of the transmitter according to the satellite and the engine parameters as follows: Δ t = Δ v · m/F; wherein m is the spacecraft mass and F is the engine thrust.
For the above implementation, referring to the schematic diagram of the RSW coordinate system shown in fig. 2, the XYZ directions in fig. 2 correspond to the perturbation thrust directions in the three directions of the track radial direction, the lateral direction and the normal direction in the RSW coordinate system, respectively. Based on this coordinate system schematic, the gaussian perturbation equation for an orbiting satellite is as follows:
wherein the radius p = a (1-e) 2 ) (ii) a r = p/(1 + ecos (f)) represents the orbital radius; f R 、F S 、F W The perturbation acceleration components along the radial direction, the transverse direction and the normal direction of the track in the track system are respectively. Based on the eccentricity e of the near-circular orbit being about 0, the simplified perturbation equation shown in equation 8 can be obtained according to the above equation.
Secondly, regarding the phase adjustment holding, the adjustment of the semi-major axis is usually considered to be realized, and the control quantity for carrying out the semi-major axis adjustment is calculated by the phase deviation and the phase deviation change rate; furthermore, the pass drift time period T is generally considered free The adjustment of phase coincidence is completed, therefore
Further, equation 9 can be obtained, where in equation 9, Δ n is the track angular velocity that needs to be changed, specifically, Δ n>0 corresponds to the negative adjustment of the long half shaft; Δ n<0 corresponds to positive adjustment of the longer half axis.
Then, for a near-circular orbit, the method comprises
Can obtain the product
Wherein Δ n =3 Δ v/a.
In conclusion, the increment of the speed required by the satellite control is shown as formula 11. After the speed increment is obtained, the ignition duration of the transmitter can be further calculated.
For the technical solution shown in fig. 1, in some possible implementations, the performing ignition according to the latitude argument, the speed increment, and the ignition duration of the ignition position to complete the trajectory control includes:
after the ignition time length is calculated, calculating the time of the satellite reaching the ignition position according to the latitude argument corresponding to the orbit control target;
and when the ignition countdown is finished, the ignition is carried out, and after the ignition time reaches the calculated ignition duration, the engine is shut down, so that the ignition is finished.
For the above implementation, after the primary phase adjustment control in the foregoing technical solution is finished, the relevant parameters of the satellite may be initialized to perform the next autonomous orbit maintenance procedure.
Based on the same inventive concept of the foregoing technical solution, referring to fig. 3, there is shown an apparatus 30 for maintaining a constellation configuration, where the apparatus 30 may be applied to any one of other satellites in a constellation except a reference satellite, and the apparatus 30 includes: a receiving section 301, a recurrence calculation section 302, an acquisition section 303, a selection section 304, and an ignition section 305; wherein, the first and the second end of the pipe are connected with each other,
the receiving portion 301 configured to receive a nominal orbit parameter and a timestamp annotated by a ground station;
the recursion calculation part 302 is configured to obtain an average nominal orbit parameter in a period from the time stamp of receiving the upper annotation to the mth passing of the set target point by a recursion calculation according to the nominal orbit parameter of the upper annotation when the satellite receives the mth passing of the set target point;
the acquiring part 303 is configured to acquire an orbit deviation amount corresponding to the mth time when the set target point is passed, according to the average nominal orbit parameter and the orbit parameter of the satellite when the mth time when the set target point is passed;
the selecting part 304 is configured to select the latitude argument, the speed increment and the ignition duration of the ignition position according to the track deviation amount and the track control target type when the number m of times of passing through the set target point is greater than a set number threshold;
the ignition portion 305 is configured to perform ignition according to the latitude argument of the ignition position, the speed increment and the ignition time length to complete the orbit control.
In some examples, the trajectory parameter includes six trajectories, and the set target point is a descent point.
In some examples, the recurrence calculation portion 302 is configured to:
nominal average orbit parameter using nominal instantaneous orbit parameter conversion based on upper notes
And equation 2 obtains the average nominal orbit parameter in the time interval delta t from the time when the upper note timestamp is received to the ith passing of the set target point
In some examples, the acquisition portion 303 is configured to:
acquiring the track deviation phase angle when the mth time passes through the set target point based on the formula 3
According to the track deviation phase angle when the mth time passes through the set target point
And the number of times the satellite passes the set target point, and obtaining the orbit deviation phase angle speed when the mth time passes the set target point based on the formula 4
And acquiring the track inclination angle deviation delta i and the track ascent point right ascension deviation delta omega when the mth time passes through the set target point according to the expressions 5 and 6 respectively.
In some examples, the pick portion 304 is configured to:
selecting a latitude argument u of the ignition position according to the type of the track control target and the formula 7;
according to the Gaussian perturbation equation of the on-orbit motion satellite in the RSW coordinate system and the orbit eccentricity e of the near-circular orbit, the simplified perturbation equation shown in the formula 8 is obtained, wherein the approximation is 0;
according to track deviation phase angle
Deviation from track phase angular velocity
And formula 9, acquiring the track angular velocity required to be changed;
obtaining equation 10 based on the setting of the near-circular orbit;
by combining formula 8 and formula 10, the velocity increment Δ v required for obtaining the satellite is shown as formula 11;
calculating and obtaining the ignition duration of the transmitter according to the satellite and the engine parameters as follows: Δ t = Δ v · m/F; wherein m is the spacecraft mass and F is the engine thrust.
In some examples, the ignition portion 305 is configured to:
after the ignition time length is calculated, calculating the time of the satellite reaching the ignition position according to the latitude argument corresponding to the orbit control target;
and when the ignition countdown is finished, the ignition is carried out, and after the ignition time reaches the calculated ignition duration, the engine is shut down, so that the ignition is finished.
It is understood that in this embodiment, "part" may be part of a circuit, part of a processor, part of a program or software, etc., and may also be a unit, and may also be a module or a non-modular.
In addition, each component in this embodiment may be integrated into one processing unit, or each unit may exist alone physically, or two or more units are integrated into one unit. The integrated unit can be realized in a form of hardware or a form of a software functional module.
Based on the understanding that the technical solution of the present embodiment essentially or a part contributing to the prior art, or all or part of the technical solution may be embodied in the form of a software product stored in a storage medium, and include several instructions for causing a computer device (which may be a personal computer, a server, or a network device, etc.) or a processor (processor) to execute all or part of the steps of the method of the present embodiment. 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.
Therefore, the present embodiment provides a computer storage medium, which stores a program for maintaining a constellation configuration, and the program for maintaining a constellation configuration implements the method steps for maintaining a constellation configuration in the above technical solution when executed by at least one processor.
Referring to fig. 4, a specific hardware structure of a satellite-borne computing device 40 capable of implementing the apparatus 30 for maintaining constellation configuration according to the above apparatus 30 for maintaining constellation configuration and a computer storage medium is shown, where the satellite-borne computing device 40 includes: a communication interface 401, a memory 402 and a processor 403; the various components are coupled together by a bus system 404. It is understood that the bus system 404 is used to enable communications among the components. The bus system 404 includes a power bus, a control bus, and a status signal bus in addition to a data bus. For clarity of illustration, however, the various buses are labeled as bus system 404 in FIG. 4. Wherein the content of the first and second substances,
the communication interface 401 is configured to receive and transmit signals during information transmission and reception with other external network elements;
the memory 402 for storing a computer program operable on the processor 403;
the processor 403 is configured to execute the steps of the method for maintaining constellation configuration in the above technical solution when the computer program is executed.
It will be appreciated that memory 402 in embodiments of the invention may be either volatile memory or nonvolatile memory, or may include both volatile and nonvolatile memory. The non-volatile Memory may be a Read-Only Memory (ROM), a Programmable ROM (PROM), an Erasable PROM (EPROM), an Electrically Erasable PROM (EEPROM), or a flash Memory. Volatile Memory can be Random Access Memory (RAM), which acts as external cache Memory. By way of illustration and not limitation, many forms of RAM are available, such as Static random access memory (Static RAM, SRAM), dynamic Random Access Memory (DRAM), synchronous Dynamic random access memory (Synchronous DRAM, SDRAM), double Data Rate Synchronous Dynamic random access memory (ddr Data Rate SDRAM, ddr SDRAM), enhanced Synchronous SDRAM (ESDRAM), synchlink DRAM (SLDRAM), and Direct Rambus RAM (DRRAM). The memory 402 of the systems and methods described herein is intended to comprise, without being limited to, these and any other suitable types of memory.
And processor 403 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 403. The Processor 403 may be a general-purpose Processor, a Digital Signal Processor (DSP), an Application Specific Integrated Circuit (ASIC), a Field Programmable Gate Array (FPGA) or other Programmable logic device, discrete Gate or transistor logic device, or discrete hardware components. The various methods, steps and logic blocks disclosed in the embodiments of the present invention may be implemented or performed. 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 402, and the processor 403 reads the information in the memory 402, and completes the steps of the method in combination with the hardware.
It is to be understood that the embodiments described herein may be implemented in hardware, software, firmware, middleware, microcode, or any combination thereof. For a hardware implementation, the Processing units may be implemented within one or more Application Specific Integrated Circuits (ASICs), digital Signal Processors (DSPs), digital Signal Processing Devices (DSPDs), programmable Logic Devices (PLDs), field Programmable Gate Arrays (FPGAs), general purpose processors, controllers, micro-controllers, microprocessors, other electronic units configured to perform the functions described herein, or a combination thereof.
For a software implementation, the techniques described herein may be implemented with modules (e.g., procedures, functions, and so on) that perform the functions described herein. The software codes may be stored in a memory and executed by a processor. The memory may be implemented within the processor or external to the processor.
It can be understood that the above-mentioned exemplary technical solutions of the apparatus 30 for maintaining a constellation configuration and the on-board computing device 40 belong to the same concept as the technical solution of the method for maintaining a constellation configuration, and therefore, for details that are not described in detail in the above-mentioned technical solutions of the apparatus 30 for maintaining a constellation configuration and the on-board computing device 40, reference may be made to the description of the technical solution of the method for maintaining a constellation configuration. The embodiments of the present invention will not be described in detail herein.
It should be noted that: the technical schemes described in the embodiments of the present invention can be combined arbitrarily without conflict.
The above description is only for the specific embodiments of the present invention, but the scope of the present invention is not limited thereto, and any person skilled in the art can easily conceive of the changes or substitutions within the technical scope of the present invention, and all the changes or substitutions should be covered within the scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the appended claims.
Claims (9)
1. A method of maintaining a constellation configuration, the method being applied to any one of the other satellites in the constellation except a reference satellite, the method comprising:
receiving nominal orbit parameters and timestamps annotated by a ground station;
when the satellite receives the mth time passing through the set target point, obtaining an average nominal orbit parameter in a time period from the time stamp of receiving the upper note to the mth time passing through the set target point through recursive calculation according to the nominal orbit parameter of the upper note;
acquiring the corresponding orbit deviation amount when the mth time passes through the set target point according to the average nominal orbit parameter and the satellite orbit parameter when the mth time passes through the set target point;
when the number m of times of passing through the set target point is greater than a set number threshold, selecting a latitude argument, a speed increment and an ignition duration of an ignition position according to the track deviation amount and the track control target type;
igniting according to the latitude argument, the speed increment and the ignition duration of the ignition position to complete the track control;
selecting a latitude argument, a speed increment and an ignition duration of an ignition position according to the track deviation amount and the track control target type, wherein the selection comprises the following steps:
selecting a latitude argument u of the ignition position according to the track control target type and the formula 6:
according to the Gaussian perturbation equation of the on-orbit motion satellite in the RSW coordinate system and the orbit eccentricity e of the near-circular orbit being approximate to 0, a simplified perturbation equation shown in the formula 7 is obtained:
wherein, F S 、F W Perturbation acceleration components in the transverse direction and the normal direction of the track under an RSW coordinate system are respectively; Δ v S 、Δv W The velocity variation components along the track transverse direction and the normal direction under the RSW coordinate system are respectively; f is a true proximal angle;
is the track angular velocity; mu is an earth gravity constant; u = ω + f;
according to track deviation phase angle
Deviation from track phase angular velocity
And equation 8, obtaining the track angular velocity Δ n that needs to be changed:
wherein, T free Representing a drift time length;
based on the setting of the near-circular orbit, equation 9 is obtained:
Δn＝3Δv/a (9)
where Δ v represents the required velocity increment for the satellite;
combining equation 7 and equation 9, the velocity increment Δ v required to obtain a satellite is shown as equation 10:
calculating and obtaining the ignition duration of the transmitter according to the satellite and the engine parameters as follows: Δ t = Δ vm/F; wherein m is the spacecraft mass and F is the engine thrust.
2. The method of claim 1, wherein the trajectory parameters include six trajectories and the set target point is a descent point.
3. The method according to claim 1, wherein the obtaining the average nominal track parameter in the period from the time stamp of the received upper note to the ith passing of the set target point by recursive calculation according to the nominal track parameter of the upper note comprises:
nominal average orbit parameter using nominal instantaneous orbit parameter conversion based on upper notes
And equation 1 obtains the average nominal orbit parameter in the time interval delta t from the time when the upper note timestamp is received to the ith passing of the set target point
In the above-mentioned formula, the compound of formula,
and
respectively represent
The first differential of the function.
4. The method according to claim 3, wherein the obtaining the orbit deviation amount corresponding to the mth time of passing through the set target point according to the average nominal orbit parameter and the orbit parameter of the satellite at the mth time of passing through the set target point comprises:
acquiring the track deviation phase angle when the mth time passes through the set target point based on the formula 2
Wherein M is m Mean anomaly, ω, representing the transformation from the measured anomaly at the mth pass through the set target point m Mean perigee argument, M, representing the transformation from the measured perigee argument at the mth pass through the set target point dm Denotes the mean nominal mean approximate point angle, ω, at the mth pass through the set target point dm Representing an average nominal perigee argument at the mth pass through the set target point;
according to the track deviation phase angle when the mth time passes through the set target point
And the number of times the satellite passes the set target point, and obtaining the orbit deviation phase angle speed when the satellite passes the set target point for the mth time based on the formula 3
Wherein, the first and the second end of the pipe are connected with each other,
represents the track deviation phase angle at the time of the 1 st pass through the set target point;
acquiring a track inclination angle deviation delta i and a track ascent intersection right ascension deviation delta omega when the mth time passes through the set target point according to equations 4 and 5:
Δi＝i m -i dm (4)
ΔΩ＝Ω m -Ω dm (5)
wherein i m Average track inclination angle, i, representing the conversion from the measured track inclination angle at the m-th pass through the set target point dm Denotes the mean nominal orbital inclination, Ω, at the m-th pass through the set target point m Mean rising point-right ascension, Ω, representing the conversion from the measured rising point-right ascension at the mth passage through the set target point dm The average nominal ascent point right ascent at the mth pass through the set target points is shown.
5. The method of claim 1, wherein said firing based on latitude argument, velocity increment, and firing duration of firing position to achieve trajectory control comprises:
after the ignition time length is calculated, calculating the time of the satellite reaching the ignition position according to the latitude argument corresponding to the orbit control target;
and when the ignition countdown is finished, the ignition is carried out, and after the ignition time reaches the calculated ignition duration, the engine is shut down, so that the ignition is finished.
6. An apparatus for maintaining a constellation configuration, the apparatus being adapted for use with any one of a plurality of satellites in a constellation other than a reference satellite, the apparatus comprising: a receiving part, a recursion calculating part, an obtaining part, a selecting part and an ignition part; wherein, the first and the second end of the pipe are connected with each other,
the receiving portion configured to receive nominal orbit parameters and timestamps annotated by a ground station;
the recursion calculation part is configured to obtain an average nominal orbit parameter in a period from the time stamp of the received upper note to the mth passing of the set target point through recursion calculation according to the nominal orbit parameter of the upper note when the satellite receives the mth passing of the set target point;
the acquisition part is configured to acquire the corresponding orbit deviation amount when the set target point is passed the mth time according to the average nominal orbit parameter and the orbit parameter of the satellite when the set target point is passed the mth time;
the selecting part is configured to select a latitude argument, a speed increment and an ignition duration of the ignition position according to the track deviation amount and the track control target type when the number m of times of passing through the set target point is greater than a set number threshold;
the ignition part is configured to ignite according to the latitude argument, the speed increment and the ignition time length of the ignition position so as to complete orbit control;
wherein the select portion is configured to:
selecting a latitude argument u of the ignition position according to the track control target type and the formula 6:
according to the Gaussian perturbation equation of the on-orbit motion satellite in the RSW coordinate system and the orbit eccentricity e of the near-circular orbit, which are approximate to 0, a simplified perturbation equation shown in the formula 7 is obtained:
wherein, F S 、F W Perturbation acceleration components along the transverse direction and the normal direction of the track under an RSW coordinate system are respectively; Δ v S 、Δv W The velocity variation components along the track transverse direction and the normal direction under the RSW coordinate system are respectively; f is a true proximal angle;
is the track angular velocity; mu is an earth gravity constant; u = ω + f;
according to track deviation phase angle
Deviation from track phase angular velocity
And equation 8, obtaining the track angular velocity Δ n that needs to be changed:
wherein, T free Representing a drift time length;
based on the setting of the near-circular orbit, equation 9 is obtained:
Δn＝3Δv/a (9)
where Δ v represents the required velocity increment for the satellite;
combining equation 7 and equation 9, the velocity increment Δ v required to obtain a satellite is shown as equation 10:
calculating and obtaining the ignition duration of the transmitter according to the satellite and the engine parameters as follows: Δ t = Δ v · m/F; wherein m is the spacecraft mass and F is the engine thrust.
7. The apparatus of claim 6, wherein the trajectory parameter comprises six trajectories, and the set target point is a descent point.
8. An on-board computing device, the on-board computing device comprising: a communication interface, a memory and a processor; the various components are coupled together by a bus system; wherein the content of the first and second substances,
the communication interface is used for receiving and sending signals in the process of receiving and sending information with other external network elements;
the memory for storing a computer program operable on the processor;
the processor, when executing the computer program, is configured to perform the steps of the method of maintaining a constellation configuration of any of claims 1 to 5.
9. A computer storage medium, characterized in that the computer storage medium stores a program for maintaining a constellation configuration, which when executed by at least one processor implements the steps of the method for maintaining a constellation configuration of any one of claims 1 to 5.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111555151.XA CN114415716B (en) | 2021-12-17 | 2021-12-17 | Method, device and medium for maintaining constellation configuration |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111555151.XA CN114415716B (en) | 2021-12-17 | 2021-12-17 | Method, device and medium for maintaining constellation configuration |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN114415716A CN114415716A (en) | 2022-04-29 |
| CN114415716B true CN114415716B (en) | 2023-02-28 |
Family
ID=81267321
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111555151.XA Active CN114415716B (en) | 2021-12-17 | 2021-12-17 | Method, device and medium for maintaining constellation configuration |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114415716B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115015983B (en) * | 2022-05-31 | 2023-03-10 | 长光卫星技术股份有限公司 | Long-term on-orbit maintaining method of optical remote sensing constellation |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO1993009613A1 (en) * | 1991-10-28 | 1993-05-13 | Calling Communications Corporation | Satellite communication system |
| CN110429974A (en) * | 2019-08-07 | 2019-11-08 | 清华大学 | Rapid alignment method and device based on regression orbit constellation |
| CN112254735A (en) * | 2020-10-15 | 2021-01-22 | 中国西安卫星测控中心 | Method for keeping constellation configuration under multiple constraint conditions |
| WO2021102669A1 (en) * | 2019-11-26 | 2021-06-03 | 中国科学院微小卫星创新研究院 | Autonomous maintenance method for ultra-low orbit satellites |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112783183B (en) * | 2020-12-29 | 2023-12-19 | 中国人民解放军63921部队 | Track planning method for sun synchronous circle regression track |
| CN113031638B (en) * | 2021-03-10 | 2022-08-23 | 哈尔滨工业大学 | Constellation deployment method, device and storage medium based on time-sharing phase modulation of low-thrust satellite |
-
2021
- 2021-12-17 CN CN202111555151.XA patent/CN114415716B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO1993009613A1 (en) * | 1991-10-28 | 1993-05-13 | Calling Communications Corporation | Satellite communication system |
| CN110429974A (en) * | 2019-08-07 | 2019-11-08 | 清华大学 | Rapid alignment method and device based on regression orbit constellation |
| WO2021102669A1 (en) * | 2019-11-26 | 2021-06-03 | 中国科学院微小卫星创新研究院 | Autonomous maintenance method for ultra-low orbit satellites |
| CN112254735A (en) * | 2020-10-15 | 2021-01-22 | 中国西安卫星测控中心 | Method for keeping constellation configuration under multiple constraint conditions |
Non-Patent Citations (7)
| Title |
|---|
| Investigating suitable orbits for the Swarm constellation mission – The frozen orbit;Kyoung-Min Roh;《sciencedirect》;20080318;全文 * |
| Model-Free Prescribed Performance Control for Spacecraft Attitude Tracking;耿云海;《IEEE TRANSACTIONS ON CONTROL SYSTEMS TECHNOLOGY》;20200224;全文 * |
| Study on Evolution Analysis and Maintenance Technology of Co-location Configuration of Geostationary Satellites;Ning Xin;《Proceedings of the 38th Chinese Control Conference》;20190730;全文 * |
| 一种考虑摄动影响的星座构型稳定性设计方法;陈长春;《上海航天》;20200225;第34-35页 * |
| 低轨Ｗａｌｋｅｒ星座构型偏置维持控制方法分析;李玖阳;《中国空间科学技术》;20210425;全文 * |
| 卫星编队飞行相对轨迹优化与控制;王有亮;《中国优秀硕博士论文全文数据库》;20190115;第2章 * |
| 阿斯图微卫星姿态控制系统设计;吴凡;《光学精密工程》;20201031;全文 * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN114415716A (en) | 2022-04-29 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| Kouba | A simplified yaw-attitude model for eclipsing GPS satellites | |
| CN108761507B (en) | Navigation satellite orbit rapid recovery method based on short arc orbit determination and prediction | |
| US9798008B2 (en) | Method of guidance for placing a satellite on station | |
| US7786929B2 (en) | Method and device for predicting GNSS satellite trajectory extension data used in mobile apparatus | |
| CN103364805A (en) | Method and system for automatically tracking medium earth orbit navigational satellite through servo antenna | |
| CN115639580A (en) | Satellite autonomous navigation method | |
| CN113031638B (en) | Constellation deployment method, device and storage medium based on time-sharing phase modulation of low-thrust satellite | |
| Tortora et al. | Precise Cassini navigation during solar conjunctions through multifrequency plasma calibrations | |
| CN114415716B (en) | Method, device and medium for maintaining constellation configuration | |
| Narayana et al. | Hummingbird: Energy efficient gps receiver for small satellites | |
| CN107300700B (en) | Agile synthetic aperture radar satellite bunching mode attitude maneuver demand calculation method | |
| Kassas et al. | Ad astra: Simultaneous tracking and navigation with megaconstellation LEO satellites | |
| CN115743619A (en) | Maneuvering and maintaining method, device and medium for satellite subsatellite point trajectory of ultra-low orbit | |
| CN109606739B (en) | Method and device for correcting earth-moon transfer orbit of detector | |
| CN111024128B (en) | Method for transmitting and aligning stable state of optical axis of airborne photoelectric pod | |
| Borisenko et al. | On the rapid orbital attitude control of manned and cargo spacecraft Soyuz MS and Progress MS | |
| CN114313313B (en) | Method, device and medium for full-electric propulsion of initial orbit distribution of small satellite to circular orbit | |
| CN115993777A (en) | Track perturbation model inversion-based diameter-cut joint control decoupling iteration calibration method | |
| CN115465475B (en) | Inverse orbit intersection detection method and device for large-scale constellation and storage medium | |
| Possner et al. | Control of science orbits about planetary satellites | |
| Kozorez et al. | Resolving the Navigation and Control Problems of Payload Insertion into a Geostationary Orbit on the Basis of the Modern and Potential GNSS Technologies | |
| Wolff et al. | Navigation considerations for low-thrust planetary missions | |
| Husain et al. | Generation of GNSS Observables for LEO Satellites Using High-Fidelity Models | |
| KR20070040085A (en) | Ground control method for collocated geostationay satellites | |
| Yoon et al. | PRECISE CALIBRATION OF MULTI-SEGMENT MANEUVERS FOR EUMETSAT POLAR SYSTEM (EPS) OPERATIONS PLANNING |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |