Aircraft orbit decay rate prediction method and device based on auroral arc charging enhancement

Abstract

This invention provides a method and apparatus for predicting the orbital decay rate of an aircraft based on aurora arc charging enhancement, belonging to the field of aircraft technology. The method uses an ionization drag correction model incorporating aurora arc charging increments to predict the historical state information of a target aircraft, obtaining the orbital decay rate of the target aircraft in the future time period. The aforementioned aurora arc charging increments can adaptively change with time, space, and environment, ensuring not only a wide applicability of the ionization drag correction model but also improving the accuracy of the predicted orbital decay rate based on the ionization drag correction model. This enables precise prediction of the aircraft's orbital position, allowing the aircraft to successfully complete its mission.

Description

Technical Field

[0001] This invention relates to the field of aircraft technology, and in particular to a method and apparatus for predicting the orbital decay rate of an aircraft based on aurora arc charging enhancement. Background Technology

[0002] During mission execution, drag (such as atmospheric drag and ionization drag) continuously consumes the spacecraft's orbital energy, causing orbital decay (i.e., shortening of the orbital semi-major axis). Timely prediction of the orbital decay rate allows for accurate prediction of the spacecraft's orbital position, enabling the spacecraft to successfully complete its mission.

[0003] However, existing spacecraft orbit decay rate prediction schemes suffer from inaccurate predictions and limited applicability, resulting in inaccurate spacecraft orbital positions predicted based on orbit decay rates, which affects the successful completion of spacecraft missions. Summary of the Invention

[0004] This invention proposes a method and device for predicting the orbital decay rate of a spacecraft based on aurora arc charging enhancement. It employs an ionization drag correction model incorporating aurora arc charging increments to predict the historical state information of a target spacecraft, thereby obtaining the orbital decay rate of the target spacecraft in the future time period. The aforementioned aurora arc charging increments can adaptively change with time, space, and environment, ensuring not only a wide applicability of the ionization drag correction model but also improving the accuracy of the predicted orbital decay rate. This enables precise prediction of the spacecraft's orbital position, allowing the spacecraft to successfully complete its mission.

[0005] To achieve the above objectives, the present invention adopts the following technical solution: Firstly, this invention provides a method for predicting the orbital decay rate of a spacecraft based on auroral arc charging enhancement. This method is used to predict the orbital decay rate of a target spacecraft during a low solar activity year due to ionization drag caused by auroral arc charging enhancement. The method includes: firstly, acquiring historical state information of the target spacecraft; the target spacecraft is a low Earth orbit spacecraft; the historical state information includes the number of two orbital elements of the target spacecraft at each time point within a historical time period. Then, based on the historical state information and an ionization drag correction model, the orbital decay rate of the target spacecraft in a future time period is predicted; the ionization drag correction model incorporates auroral arc charging increments. The ionization resistance model for aurora arc charging enhancement was modified; aurora arc charging increment. Determined by geomagnetic coordinates and ionization environment parameters; geomagnetic coordinates are used to define the spatiotemporal extent of auroral arc charging, while ionization environment parameters indicate solar activity to adjust the auroral arc charging increment. Size.

[0006] The method for predicting the orbital decay rate of a spacecraft based on auroral arc charging enhancement provided by this invention first acquires the historical state information of the target spacecraft, which includes the number of two orbital elements at each time point within multiple time periods in the historical time period. Then, an ionization drag correction model that incorporates auroral arc charging increments to correct for auroral arc charging enhancement is used to predict the orbital decay rate of the target spacecraft in the future time period. In the above process, the auroral arc charging increment is defined by the geomagnetic coordinates at the time and space of auroral arc charging and by the ionization environment parameters, so that the auroral arc charging increment determined by the geomagnetic coordinates and ionization environment parameters can adaptively change with time, space, and environment. This not only ensures the wide applicability of the ionization drag correction model but also improves the prediction accuracy of the spacecraft orbital decay rate predicted based on the ionization drag correction model, thereby achieving accurate prediction of the spacecraft's orbital position and enabling the spacecraft to successfully complete its mission.

[0007] In one implementation of the first aspect, based on historical state information and an ionization drag correction model, the orbital decay rate of a target spacecraft over a future time period is predicted. This includes: for each of multiple time points, firstly, inputting the two orbital elements, ionization environment parameters, and geomagnetic coordinates of the target spacecraft at that time point into the ionization drag correction model to calculate the ionization drag experienced by the target spacecraft at that time point. Then, based on the ionization drag experienced by the target spacecraft at that time point, the orbital decay rate of the target spacecraft at that time point is calculated. Finally, the orbital decay rates of the target spacecraft at multiple time points are integrated to obtain the orbital decay rate of the target spacecraft over the future time period.

[0008] In one implementation of the first aspect, the orbital decay rate of the target spacecraft at a given time point is calculated based on the ionizing drag experienced by the target spacecraft at that time point. This includes: first, dividing the ionizing drag experienced by the target spacecraft at that time point by the mass of the target spacecraft to obtain the ionizing drag acceleration of the target spacecraft at that time point; then, using the ionizing drag acceleration of the target spacecraft at that time point, the orbital decay rate of the target spacecraft at that time point is calculated; the formula for calculating the orbital decay rate is: ,in, The rate of change of the semi-major axis of the orbit. Indicates the orbital decay rate; For the semi-major axis of the track; The gravitational constant of the central celestial body; i Indicates the ionization resistance indicator. This represents the ionization drag acceleration of a target aircraft distributed tangentially.

[0009] In one implementation of the first aspect, obtaining the historical state information of the target spacecraft includes: first, obtaining the set of two rows of orbital elements corresponding to the available epochs of the target spacecraft within a historical time period; then, using a J2 perturbation model to densify the set of two rows of orbital elements to obtain the historical state information of the target spacecraft; the time difference between any two consecutive time points among multiple time points is less than the time difference between any two consecutive available epochs of the target spacecraft.

[0010] In one implementation of the first aspect, the method further includes: optimizing the ionization resistance correction model.

[0011] In one implementation of the first aspect, the ionization environment parameters include electron number density, ion number density, and solar activity indicators.

[0012] Secondly, this invention provides a spacecraft orbit decay rate prediction device based on auroral arc charging enhancement, used to predict the orbit decay rate of a target spacecraft during a low solar activity year due to ionization drag caused by auroral arc charging enhancement. The device includes an acquisition module and a prediction module. The acquisition module is used to acquire historical state information of the target spacecraft; the target spacecraft is a low Earth orbit spacecraft; the historical state information includes the number of two orbital elements of the target spacecraft at each time point within multiple time points in a historical period. The prediction module is used to predict the orbit decay rate of the target spacecraft in a future time period based on the historical state information and an ionization drag correction model; the ionization drag correction model incorporates auroral arc charging increments. The ionization resistance model for modifying the aurora arc charging enhancement effect; aurora arc charging increment It is determined by geomagnetic coordinates and ionization environment parameters; the ionization environment parameters include electron number density, ion number density, and solar activity indicators.

[0013] In one implementation of the first and second aspects, the aurora arc charging increment satisfy: Represents the magnitude function, The solar activity index F10.7 indicates... Indicates the basic amplitude. Indicates the increase in amplitude. Indicates the first activation slope. Indicates the first activation boundary. Represents the power exponent; magnitude function satisfy: Represents the magnetic latitude activation function. Indicates magnetic latitude. Indicates the second activation slope. Indicates the second activation boundary; magnetic latitude activation function satisfy: This represents the local activation function for the magnetic field. When indicating a magnetic place, Indicates the peak value at local magnetic field time. Indicates the time span of activation of the aurora charging phenomenon; magnetic local time activation function. satisfy: This represents the difference between the local magnetic time and the peak value. Indicates the local time peak value of the magnetic field.

[0014] In one implementation of the first and second aspects, the ionization resistance correction model satisfies: in, i Indicates the ionization resistance indicator. This represents the ionization drag experienced by the target aircraft; This represents the ion number density of the target aircraft; Indicates the radius of the target aircraft; Represents a unit charge; Indicates the total charging amount; Indicates the average mass of ions in the environment; This indicates the relative velocity of the target aircraft; Indicates the length of the Debye shield; Indicates impact parameters; total charge. satisfy , This represents the surface charging equilibrium potential of the target aircraft. This indicates the incremental charge of the aurora arc.

[0015] Thirdly, the present invention provides an electronic device including a processor and a memory coupled to the processor; the memory is used to store computer instructions, and when the electronic device is running, the processor executes the computer instructions stored in the memory to cause the electronic device to perform the method as described in the first aspect above or any implementation thereof.

[0016] Fourthly, the present invention provides a computer-readable storage medium including computer program instructions that, when executed by a computer, cause the computer to perform the method described in the first aspect above or any implementation thereof.

[0017] Fifthly, the present invention provides a computer program product, including computer program instructions, which, when executed on a computer, cause the computer to perform the method described in the first aspect above or any implementation thereof.

[0018] The technical effects corresponding to the second to fifth aspects and their possible implementations can be referred to the above description of the technical effects of the first aspect and its possible implementations, and will not be repeated here. Attached Figure Description

[0019] Figure 1 This is a schematic diagram showing the superposition of data on the orbital semi-major axis, normalized tangential acceleration, and F10.7 exponent of the formation constellation on the same time axis, as provided in the embodiments of this application. Figure 2 This is a schematic diagram showing the superposition of ion and electron number densities on the same time axis in the formation constellation provided in this application embodiment; Figure 3 This is a schematic diagram comparing the changing trends of the tangential acceleration and the F10.7 exponent of the formation constellation provided in this application embodiment; Figure 4 This is one of the schematic diagrams of the spacecraft orbit decay rate prediction method based on aurora arc charging enhancement provided in the embodiments of this application; Figure 5 This is the second schematic diagram of the method for predicting the orbital decay rate of a spacecraft based on aurora arc charging enhancement provided in the embodiments of this application; Figure 6 This is the third schematic diagram of the method for predicting the orbital decay rate of a spacecraft based on aurora arc charging enhancement provided in the embodiments of this application; Figure 7 This is a schematic diagram of the distribution of remote sensing satellite trajectories in magnetic latitude and magnetic local time provided in the embodiments of this application; Figure 8 This is a schematic diagram comparing the remote sensing satellite trajectory in latitude and magnetic latitude according to an embodiment of this application; Figure 9 This is a schematic diagram of the aurora charging event distribution described by the magnetic latitude activation function and the magnetic local time activation function provided in the embodiments of this application; Figure 10 This is the fourth schematic diagram of the spacecraft orbit decay rate prediction method based on aurora arc charging enhancement provided in the embodiments of this application; Figure 11 This is a schematic diagram of the structure of the spacecraft orbit decay rate prediction device based on aurora arc charging enhancement provided in the embodiments of this application. Detailed Implementation

[0020] In the specification and claims of this invention, the terms "first" and "second," etc., are used to distinguish different objects, rather than to describe a specific order of objects.

[0021] In the embodiments of this application, the terms "exemplary" or "for example" are used to indicate that something is an example, illustration, or description. Any embodiment or design that is described as "exemplary" or "for example" in the embodiments of this application should not be construed as being more preferred or advantageous than other embodiments or design. Specifically, the use of the terms "exemplary" or "for example" is intended to present the relevant concepts in a specific manner.

[0022] In the description of this invention, unless otherwise stated, "a plurality of" means two or more. For example, "a plurality of time points" means two or more time points.

[0023] The method and apparatus provided in this application relate to predicting the orbital decay rate of a spacecraft, and can be used to predict the orbital decay rate of a target spacecraft affected by the enhanced ionization drag caused by auroral arc charging during a low solar activity year.

[0024] As is understandable, solar activity refers to the collective term for various violent dynamic phenomena occurring in the Sun's atmosphere, mainly manifested as sunspots, flares, prominences, and solar wind. These phenomena originate from drastic changes in the Sun's magnetic field and typically occur in the photosphere, chromosphere, and corona.

[0025] The solar activity cycle (approximately eleven years) refers to the quasi-periodic changes caused by the activity of the solar magnetic field. High solar activity years and low solar activity years are two typical phases of the solar activity cycle. High solar activity years are those with a large number of sunspots and intense solar activity, typically occurring at the peak of the solar activity cycle; low solar activity years are those with very few sunspots and quiet solar activity, occurring at the trough of the solar activity cycle.

[0026] Low Earth Orbit (LEO) spacecraft / constellations of LEO spacecraft, during mission execution, need to orbit a central celestial body (such as Earth) within a predetermined orbit (such as the LEO region) to successfully complete their missions. To ensure the operational stability of LEO spacecraft, accurate prediction of their orbital positions is necessary. However, LEO spacecraft experience drag during operation, causing orbital decay (shortening of the orbital semi-major axis).

[0027] Taking Earth as the central celestial body as an example, among the dissipative perturbations that cause orbital decay and shape changes, the most significant influence comes from the drag caused by the residual atmosphere or plasma in space. Most LEO spacecraft operate in the high-density ionosphere (100–1500 km). In addition to the atmospheric drag from the background residual atmosphere, the charging effect on the spacecraft surface can also generate negligible ionizing drag through particle collection and deflection. This charging effect refers to the process by which, after electrons and ions collide with the spacecraft surface, the spacecraft emits electrons outward due to processes such as the photoelectric effect. When the two currents are unbalanced, a voltage relative to the surrounding plasma is established on the spacecraft surface.

[0028] Ionizing drag can be interpreted as the drag that reduces the speed of a spacecraft caused by Coulomb scattering and hyperbolic trajectories as charged particles escape from the electric field region of the charging satellite on the spacecraft surface. It should be understood that although the density of ionospheric plasma is much lower than that of neutral gas, the ionizing drag mechanism and the expansion of the sheath (meaning the increased thickness of the plasma sheath surrounding the spacecraft due to charging, resulting in an effective cross-sectional area for interaction between the spacecraft and the plasma that is much larger than its own geometric area) will generate a drag intensity exceeding that of atmospheric drag in the upper LEO region. Typically, the equilibrium potential of a low-Earth orbit (LEO) spacecraft is approximately -1.5V, and the corresponding ionizing drag is much smaller than atmospheric drag. However, when traversing the intense auroral electron flux that generates auroral arcs, LEO spacecraft can be charged to a negative potential of 100V to 1000V (i.e., auroral arc charging), making the ionizing drag corresponding to the equilibrium potential a crucial factor in orbit prediction. Therefore, neglecting the proportion of ionizing drag in dissipative drag will lead to incorrect estimations of the spacecraft's orbital position.

[0029] Regarding the aforementioned ionization drag, one existing method calculates it using an ionization drag formula identical in form to atmospheric drag. This method only incorporates the Coulomb effect into the ionization drag coefficient, deriving a theoretical analysis based on the kinetic energy decay caused by the Coulomb effect. Therefore, it cannot address the increased ionization drag due to solar activity fluctuations, leading to discrepancies between predicted and actual trajectories. Another existing method uses the particle-in-cell (PIC) method to simulate charged aerodynamics under the Vlasov-Poisson equations, then uses plasma scaling coefficients to establish a surrogate model to describe ionization drag, summarizing the ionization drag patterns shown in the PIC simulation data. While this method contributes to the drag correction by incorporating the influence of space environment changes, it lacks further physical explanation and cannot be adapted to cross-orbital environments. Therefore, existing methods for determining spacecraft ionization drag suffer from inaccuracies and narrow applicability.

[0030] To address the issues of inaccurate prediction and limited applicability of existing spacecraft orbit decay rate prediction schemes, which result in inaccurate predicted orbital positions and hinder mission success, this application provides a method and apparatus for predicting spacecraft orbit decay rates based on aurora arc charging enhancement. This method employs an ionization drag correction model incorporating aurora arc charging increments to predict the historical state information of the target spacecraft, obtaining the orbit decay rate for the target spacecraft in the future time period. The aurora arc charging increments can adaptively change with time, space, and environment, thus broadening the applicability and improving the accuracy of orbit decay rate predictions based on the ionization drag correction model. This enables precise prediction of the spacecraft's orbital position, allowing the spacecraft to successfully complete its mission.

[0031] For example, the spacecraft orbital decay rate prediction method based on aurora arc charging enhancement provided in this embodiment of the invention can be executed by an electronic device with processing capabilities, such as a computer or server. Taking a computer as an example, the hardware components of the computer may include: a processor, memory, a network interface, a user interface, a communication bus, etc.

[0032] The processor controls the electronic equipment to perform related processing and computational tasks. For example, it is used to acquire historical state information of the target spacecraft, predict the target spacecraft's orbital decay rate, and optimize the ionization drag correction model. The processor may include a central processing unit (CPU) or other processors, and can be single-core or multi-core; for example, the processor may include multiple CPUs.

[0033] Memory is used to store computer instructions and related data. For example, it may be used to store historical state information of a target aircraft, ionization drag correction models, and the target aircraft's orbital decay rate. Memory can be random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM), flash memory, optical storage, disk storage media, or other magnetic storage devices, or any other medium capable of storing program code or data accessible by a computer. Optionally, memory can be integrated into the processor, or it can be independent of the processor.

[0034] A network interface is used for communication between a computer and other devices or communication networks. A network interface can be a transceiver with transmit and receive capabilities. Optionally, a network interface may include standard wired interfaces or wireless interfaces (such as Wi-Fi interfaces, Bluetooth interfaces, and 5G interfaces).

[0035] The communication bus is used to enable communication between different components. For example, the processor, memory, network interface and user interface mentioned above can be interconnected through the communication bus.

[0036] The user interface may include a display screen and an input unit (such as a keyboard). Optionally, the user interface may also include a standard wired interface or a wireless interface.

[0037] Those skilled in the art will understand that the computer described above may include more or fewer components, or combine certain components, or have different component arrangements; the embodiments of this application do not limit this.

[0038] To better understand the method and apparatus for predicting the orbital decay rate of a spacecraft based on aurora arc charging enhancement provided in the embodiments of this application, the following analysis will take the three-star formation of the remote sensing constellation group 8 as an example to analyze the negative correlation of orbital decay during the low solar activity year.

[0039] Step 1: Analyze the correlation between the tangential acceleration of the spacecraft and the environmental particle number density and the solar activity F10.7 index.

[0040] Using the Time-Like (TLE) data from eight satellite constellations (RS-9, 16, 17, 20, 25, 31-2, 31-3, 31-4), and after standard dataset processing including orbital control segment removal, outlier removal, and surface-to-mass ratio normalization (the ratio of an aircraft's effective cross-sectional area to its mass), a TLE dataset for 24 satellites was obtained. Furthermore, the TLE information was converted into Kepler orbital parameters, and after coordinate transformation to longitude, latitude, and altitude in the geodetic coordinate system, environmental ionospheric parameters such as electron number density, ion number density, and temperature were extracted using the International Reference Ionosphere (IRI) model to establish the TLE-IRI dataset.

[0041] Eight constellation groups at the same orbital altitude but with different phases or plane angles were aligned on the timeline to study their systematic dynamic responses to disturbances from solar and geomagnetic activity. Based on the TLE-IRI dataset, the orbital semi-major axis, normalized tangential acceleration, corresponding F10.7 exponent (i.e., the intensity of radio radiation emitted by the Sun at a wavelength of 10.7 cm, measured in solar flux sfu), and electron and ion number densities of all satellites are displayed on the timeline as follows: Figure 1 and Figure 2 As shown.

[0042] According to the NRLMSISE-00 model (an empirical model describing the temperature and density of the Earth's atmosphere), the average atmospheric density at an altitude of 1000 km and a latitude of 30° is approximately 3.65 × 10⁻⁶. -18 kg / m 3 Substituting the atmospheric density into the atmospheric drag formula, we can calculate the atmospheric drag and estimate that the tangential drag acceleration under typical conditions at this altitude is approximately 2.45 × 10⁻⁶. -12 N / kg, its order of magnitude is significantly smaller than that of the above. Figure 1 and Figure 2 The inversion of TLE shown f t Then it is considered that the above f t It is almost entirely contributed by ionization resistance.

[0043] The atmospheric drag formula is shown below.

[0044] in, C D This is the atmospheric drag coefficient. ρ N For neutral particle density, For the speed of the aircraft, S cross For the cross-sectional area of ​​the aircraft, m For the mass of the aircraft.

[0045] Although there is a positive correlation between environmental particle number density (the sum of electron and ion number densities) and the solar activity F10.7 index, both reach their maximum near the activity peaks of the 24- and 25-solar cycles, with the peak being higher in the 25-solar cycle. However, Figure 1 The tangential acceleration shown f t The overall evolution of [atmosphere density] shows a trend opposite to that of F10.7 and number density. This phenomenon contradicts the common understanding that "the greater the atmospheric density, the greater the drag."

[0046] Step 2: Analyze the impact of solar activity on orbital decay behavior under ionization drag.

[0047] The normalized tangential accelerations of all satellites were merged to construct a global time series. The series was then subjected to sliding statistics with a time window of 20, and the mean and standard deviation within each time window were calculated. Figure 3 The evolution trend of the normalized tangential acceleration mean and ±1 are given. σThe fluctuation range is shown, and the F10.7 index at the corresponding time and position is simultaneously given on the right axis.

[0048] Depend on Figure 3 It is known that solar activity is one of the dominant factors in ionization drag. During years of high solar activity (F10.7 ≥ 100 sfu), the ionization drag acceleration generally shows a positive correlation with the F10.7 exponent due to the significantly increased particle number density. However, during years of low solar activity (F10.7 < 100 sfu), the normalized tangential force reflected by the ionization drag acceleration exhibits an anomalous increase. This suggests that even when solar activity is weak, other mechanisms or environmental conditions still play a major role in enhancing ionization drag.

[0049] As is generally known in the field, the state of the space ionosphere is primarily modulated by both solar and geomagnetic activity. Solar activity leads to increased EUV (extreme ultraviolet) intensity and accelerated photoionization rates, resulting in a significant increase in ion number density at all altitudes globally. Meanwhile, geomagnetic activity injects a large amount of energy, causing an overall increase in density at orbital altitudes and the injection of additional particles, further increasing the number density.

[0050] However, the physical mechanisms of ionization drag interact with the environment in a more complex way. During years of high solar activity, strong solar EUV radiation causes the ionosphere to heat up, expand, and increase in density. This high-density plasma provides a more efficient charge conduction path for the spacecraft surface, helping accumulated static charge to leak out more quickly, thus inhibiting the spacecraft's charging process and affecting ionization drag. Conversely, during years of low solar activity, the ionospheric electron density decreases, and conductivity drops, inhibiting charge leakage pathways and making it easier for the spacecraft to accumulate high potential in the plasma environment. This electric field-particle interaction mechanism, caused by Coulomb's effect during years of low solar activity, may also create a significant ionization drag enhancement effect in the auroral region.

[0051] Auroral arc charging refers to the phenomenon where spacecraft in inclined or polar orbits below 2000 km are charged to a negative potential of 100 to 1000 volts when passing through the intense auroral electron flux that generates the auroral arc. Preliminary empirical evidence for auroral arc charging has been obtained from on-orbit observations by several spacecraft, including the DMSP, Freja, and the International Space Station. Analysis of charging events in the DMSP between 1989 and 2000, within solar cycles 22 to 23, shows an inverse correlation between auroral charging and solar activity. The fundamental reason for this is that a sufficiently low background plasma density is a prerequisite for generating a strong negative potential; a high background density enhances the current neutralization process, thus suppressing surface charge accumulation.

[0052] In other words, low solar activity and low background ion density together constitute the key triggering conditions for auroral arc charging and its associated additional drag effects. Therefore, introducing an ionization drag correction model enhanced by auroral arc charging can accurately predict orbital decay rates during low solar activity years.

[0053] The following provides a detailed description of the method and apparatus for predicting the orbital decay rate of a spacecraft based on aurora arc charging enhancement, as provided in the embodiments of this application.

[0054] The method for predicting the orbital decay rate of a spacecraft based on auroral arc charging enhancement provided in this application embodiment is used to predict the orbital decay rate of a target spacecraft in a low solar activity year due to ionization drag enhanced by auroral arc charging. For example, Figure 4 As shown, the method provided in this application embodiment includes S1-S2.

[0055] S1. Obtain the historical status information of the target aircraft.

[0056] The target spacecraft mentioned above is a low-Earth orbit spacecraft (e.g., a low-Earth orbit satellite). The historical status information mentioned above may include the number of two orbital elements (hereinafter referred to as TLE) of the target spacecraft at each time point within multiple time points in a historical period. Since the number of two orbital elements is a commonly used term in this technical field, this application embodiment will not further elaborate on the number of two orbital elements.

[0057] In one application scenario, combined with Figure 4 ,like Figure 5 As shown, S1 includes S101-S102.

[0058] S101. Obtain the set of two rows of orbital elements corresponding to the available epochs of the target spacecraft within the historical time period.

[0059] S102. The J2-term perturbation model is used to densify the two sets of orbital elements to obtain the historical state information of the target aircraft.

[0060] It is understandable that the time difference between any two consecutive time points in the historical state information of the target aircraft obtained after the above densification process is less than the time difference between any two consecutive available epochs of the target aircraft.

[0061] In one implementation of the above application scenario, taking the target spacecraft as a low-Earth orbit satellite as an example, let the available TLE set for low-Earth orbit satellites be... Its corresponding time epoch is Because each pair of consecutive epochs The corresponding TLE ( r i,TLE , ri+1,TLE The lack of dense state parameters between these pairs of epochs necessitates obtaining information about each pair of consecutive epochs. The orbital decay rate, in each consecutive epoch The corresponding TLE ( r i,TLE , r i+1,TLE Between these points, the state points of the dense intermediate interpolated trajectories are generated analytically and recursively using the J2-term perturbation model. .

[0062] Specifically, for a remote sensing constellation at an altitude of 1000 km, the main sources of orbital perturbations are Earth's oblateness and dissipative drag. The influence of third-object perturbations and radiation pressure perturbations is relatively low. Considering the relatively short recursive time window between consecutive TLEs, the J2-term perturbation model is sufficient to accurately capture the main evolutionary characteristics of the orbit. The state points obtained through the J2-term perturbation model... The accuracy requirements of the ionization resistance correction model for TLE are met. Since the J2 perturbation model is a commonly used technique in this field, the above-mentioned densification process will not be elaborated upon in this application's embodiments.

[0063] S2. Based on historical state information and ionization drag correction model, predict the orbital decay rate of the target spacecraft in the future time period.

[0064] In some embodiments, combined with Figure 5 ,like Figure 6 As shown, S2 above includes S201-S203.

[0065] S201. For each of the multiple time points, input the two orbital elements of the target spacecraft at that time point, the ionization environment parameters, and the geomagnetic coordinates into the ionization drag correction model to calculate the ionization drag experienced by the target spacecraft at that time point.

[0066] In some application scenarios, the ionization environment parameters and geomagnetic coordinates of the target spacecraft at the aforementioned time point can be obtained as follows: The TLE of the target spacecraft at that time point is converted into Kepler orbital elements, and then the Kepler orbital elements are converted to both the geodetic and geomagnetic coordinate systems to obtain the geodetic coordinates (longitude, latitude, and altitude) and the geomagnetic coordinates (magnetic latitude). and magnetic local time Then, the geodetic coordinates are input into the IRI model to obtain the ionization environment parameters. These ionization environment parameters may include electron number density, ion number density, and solar activity indicators, etc.

[0067] In this embodiment, the above-mentioned ionization resistance correction model incorporates the aurora arc charging increment. The ionization resistance model for enhanced aurora arc charging is modified accordingly. The aforementioned aurora arc charging increment... Determined by geomagnetic coordinates and ionization environment parameters; the geomagnetic coordinates are used to define the spatiotemporal extent of auroral arc charging, and the ionization environment parameters indicate solar activity to adjust the auroral arc charging increment. Size.

[0068] In one implementation, the above-mentioned ionization resistance correction model satisfies the following formula.

[0069] in, i Indicates the ionization resistance indicator. This represents the ionization drag experienced by the target aircraft; This represents the ion number density of the target aircraft; Indicates the radius of the target aircraft; Represents a unit charge; Indicates the total charging amount; Indicates the average ion mass in the environment; This indicates the relative velocity of the target aircraft. Obtained via TLE; The simplified formula for calculating the Debye shielding length is as follows: ; Indicates the impact parameter. It is related to the surface area of ​​the aircraft.

[0070] The above total charging amount satisfy , The surface charging equilibrium potential of the target aircraft is represented by... Calculated; Indicates the increment of aurora arc charging. It satisfies the following formula.

[0071] in, Represents the magnitude function; The solar activity index F10.7 indicates... This represents the baseline amplitude, used to fit the difference between theoretical and actual ionization drag under high solar activity. Indicates the increase in amplitude. Indicates the first activation slope. Indicates the first activation boundary. That is, the first activation boundary The transition slope of the curve at that point; It represents the power exponent. Represents the magnetic latitude activation function; Indicates the second activation slope. Indicates the second activation boundary; That is, the second activation boundary The transition slope of the curve at that point. This represents the local activation function for the magnetic field. Indicates the peak value at local magnetic field time. This indicates the duration of the aurora charging phenomenon.

[0072] The above amplitude function It is constructed from ionization environment parameters and used to modulate the intensity of solar activity to adjust the auroral arc charge increment. The magnitude of the amplitude function is modeled as a sigmoid function with respect to the F10.7 exponent, given the obvious threshold or saturation characteristics of auroral arc charging conditions. Furthermore, based on the relationship between auroral arc charging and background plasma density, the magnitude of the amplitude function is further adjusted according to the relationship between ion density and background plasma density. N i inverse proportional term Then the above amplitude function It satisfies the following formula.

[0073] Furthermore, since key driving factors such as auroral electron deposition, field current, and polar ionospheric conductivity are typically organized along geomagnetic field lines and form regions related to magnetic latitude in the geomagnetic pole regions, auroral arc charging events are mainly distributed in orbital arcs with a MLAT (magnetic latitude) greater than 60°, and between 18:00 at night and 2:00 at dawn (magnetic local time), with the highest concentration around 22:00. By monitoring over a two-day period, the distribution of a remote sensing satellite's trajectory along the magnetic latitude-magnetic local time can be obtained as follows: Figure 7 As shown, orbital arcs exist in both the Northern and Southern Hemispheres within this region. The comparison between latitude and magnetic latitude (the degree of position relative to the Earth's magnetic equator, determined by the Earth's magnetic dipole axis) is as follows: Figure 8 As shown, magnetic latitude has asymmetry between the northern and southern hemispheres and TLE data is sparse. The historical state information obtained after the above-mentioned S101 to S102 densification processing is meaningful and can meet the accuracy requirements.

[0074] The above magnetic latitude activation function magnetic latitude in geomagnetic coordinates The constructed region (corresponding to space) is used to define the area where auroral arc charging occurs. Since auroral arc charging has distinct triggering characteristics, a sigmoid function is used to describe the magnetic latitude activation function to represent the variation process from low to high latitudes. The aforementioned magnetic latitude activation function... It satisfies the following formula.

[0075] The above magnetic local time activation function Magnetic local time in geomagnetic coordinates The constructed timeframe is used to define the period (corresponding time) during which auroral arc charging occurs, based on... MLT The statistics will MLT =22h is set as the distribution peak, and most events are distributed between 22h ± 4h. A Gaussian function is used to describe the magnetic local time activation function to represent the decrease in magnetic local time on both sides of the nighttime peak. The above magnetic local time activation function... It satisfies the following formula.

[0076] This represents the difference between the local magnetic time and the peak value. Indicates the local time peak value of the magnetic field.

[0077] The above magnetic latitude activation function And the above magnetic local time activation function The physical intuition summarized from the measured data of the aircraft is transformed into mathematical formulas, and its distribution and activation level are as follows: Figure 9 As shown. By Figure 9 It can be seen that the above activation function fits... MLAT - MLT The distribution results are consistent with the aurora charging event statistics of Freja and DMSP satellites.

[0078] S202. Calculate the orbital decay rate of the target spacecraft at the time point based on the ionization drag experienced by the target spacecraft at that time point.

[0079] In one embodiment, combined with Figure 6 ,like Figure 10 As shown, S202 above includes S2021-S2022.

[0080] S2021. Divide the ionization drag experienced by the target spacecraft at the time point by the mass of the target spacecraft to obtain the ionization drag acceleration of the target spacecraft at the time point.

[0081] Specifically, the formula for calculating the ionization drag acceleration is as follows.

[0082] in, This represents the ionization drag acceleration of a target aircraft distributed tangentially. Indicates the resistivity coefficient, and ; This represents the cross-sectional area of ​​the aircraft.

[0083] S2022. Using the ionization drag acceleration of the time-point target spacecraft, calculate the orbital decay rate of the time-point target spacecraft.

[0084] Optionally, the formula for calculating the above-mentioned orbital attenuation rate is: ,in, The rate of change of the semi-major axis of the orbit. Indicates the orbital decay rate; For the semi-major axis of the track; is the gravitational constant of the central celestial body.

[0085] S203. Integrate the orbital decay rate of the target spacecraft at multiple time points to obtain the orbital decay rate of the target spacecraft in the future time period.

[0086] It should be noted that since the above-mentioned integration of the orbital decay rate is a commonly used technical means in this field, the embodiments of this application will not be described in detail here.

[0087] In some application scenarios, the above method also includes S3.

[0088] S3. Optimize the ionization resistance correction model.

[0089] In one implementation, a genetic algorithm is used to optimize the ionization drag correction model to obtain an optimized ionization drag correction model. For example, the optimization objective is to minimize the average time deviation between the predicted and measured values ​​of the orbital decay rate, and the parameters of the ionization drag correction model are adjusted accordingly. Optimization was performed to obtain the optimized ionization resistance correction model parameters. Since using genetic algorithms to optimize model parameters is a common technique in this field, the optimization process described above will not be elaborated upon in the embodiments of this application.

[0090] It should be noted that S3 can be executed after S2 or before S2. This application embodiment does not limit the execution order of S3.

[0091] In summary, the method for predicting the orbital decay rate of a spacecraft based on auroral arc charging enhancement provided in this application first obtains the historical state information of the target spacecraft, which includes the number of two orbital elements at each time point within multiple time points in a historical time period. Then, it uses an ionization drag correction model that incorporates auroral arc charging increments to correct for auroral arc charging enhancement to predict the orbital decay rate of the target spacecraft in the future time period. In the above process, the auroral arc charging increment is limited by the geomagnetic coordinates to determine the spatiotemporal occurrence of auroral arc charging and by the ionization environment parameters to limit its value. This allows the auroral arc charging increment determined by the geomagnetic coordinates and ionization environment parameters to adapt to changes in time, space, and environment. This not only ensures a wide range of applications for the ionization drag correction model but also improves the prediction accuracy of the spacecraft orbital decay rate predicted based on the ionization drag correction model, thereby achieving accurate prediction of the spacecraft's orbital position and enabling the spacecraft to successfully complete its mission.

[0092] Accordingly, this application provides a spacecraft orbital decay rate prediction device based on auroral arc charging enhancement, used to predict the orbital decay rate of a target spacecraft in a low solar activity year due to ionization drag enhanced by auroral arc charging. Figure 11 As shown, the above-mentioned device includes an acquisition module 501 and a prediction module 502.

[0093] The acquisition module 501 is used to acquire historical state information of the target spacecraft; the target spacecraft is a low-Earth orbit spacecraft; the historical state information includes the number of two orbital elements of the target spacecraft at each time point within multiple time points in a historical time period. For example, the acquisition module 501 is used to implement S1 of the above method.

[0094] Prediction module 502 is used to predict the orbital decay rate of a target spacecraft over a future time period based on historical state information and an ionization drag correction model; the ionization drag correction model incorporates the aurora arc charging increment. The ionization resistance model for modifying the aurora arc charging enhancement effect; aurora arc charging increment It is determined by geomagnetic coordinates and ionization environment parameters; the ionization environment parameters include electron number density, ion number density, and solar activity indicators. For example, prediction module 502 is used to implement S2 of the above method.

[0095] Optionally, the prediction module 502 is specifically used to: for each of multiple time points, input the two orbital elements, ionization environment parameters, and geomagnetic coordinates of the target spacecraft at that time point into the ionization drag correction model to calculate the ionization drag experienced by the target spacecraft at that time point. Based on the ionization drag experienced by the target spacecraft at that time point, calculate the orbital decay rate of the target spacecraft at that time point. Integrate the orbital decay rates of the target spacecraft at multiple time points to obtain the orbital decay rate of the target spacecraft in the future time period. For example, the prediction module 502 is specifically used to implement S201-S203 of the above method.

[0096] Preferably, the above-mentioned device further includes an optimization module 503.

[0097] The optimization module 503 is used to optimize the ionization resistance correction model. For example, the optimization module 503 is used to implement S3 of the above method.

[0098] The modules of the above-mentioned spacecraft orbit decay rate prediction device based on aurora arc charging enhancement can also be used to perform other steps in the above method embodiments. All relevant contents involved in the above method embodiments can be referred to the functional descriptions of the corresponding functional modules, and will not be repeated here.

[0099] This application also provides an electronic device, including: a processor and a memory coupled to the processor; the memory is used to store computer instructions, and when the electronic device is running, the processor executes the computer instructions stored in the memory to cause the electronic device to perform the methods in the above embodiments. The processor can implement the acquisition module 501, the prediction module 502, and the optimization module 503 described above; the memory can also be used to store historical state information of the target spacecraft, an ionization drag correction model, and the target spacecraft's orbital decay rate, etc.

[0100] This application also provides a computer-readable storage medium including a computer program that, when run on a computer, performs the methods described in the above embodiments.

[0101] This application also provides a computer program product, which includes computer program instructions that, when run on a computer, execute the methods described in the above embodiments.

[0102] The various embodiments in this specification are described in a progressive manner. The same or similar parts between the various embodiments can be referred to each other. Each embodiment focuses on describing the differences from other embodiments.

[0103] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of this application.

Claims

1. A method for predicting the orbital decay rate of a spacecraft based on aurora arc charging enhancement, characterized in that, The method for predicting the orbital decay rate of a target spacecraft affected by ionization drag enhanced by auroral arc charging during a low solar activity year includes: Obtain the historical state information of the target spacecraft; the target spacecraft is a low Earth orbit spacecraft; the historical state information includes the number of two orbital elements of the target spacecraft at each time point within multiple time points in a historical period; Based on the historical state information and the ionization drag correction model, the orbital decay rate of the target spacecraft in the future time period is predicted; the ionization drag correction model incorporates the aurora arc charging increment. The ionization resistance model for aurora arc charging enhancement is modified; the aurora arc charging increment Determined by geomagnetic coordinates and ionization environment parameters; the geomagnetic coordinates are used to define the spatiotemporal extent of auroral arc charging, and the ionization environment parameters indicate solar activity to adjust the auroral arc charging increment. Size.

2. The method as described in claim 1, characterized in that, The aurora arc charging increment satisfy: Represents the magnitude function, The solar activity index F10.7 indicates... Indicates the basic amplitude. Indicates the increase in amplitude. Indicates the first activation slope. Indicates the first activation boundary. The magnitude function represents the power exponent. satisfy: Represents the magnetic latitude activation function. Indicates magnetic latitude. Indicates the second activation slope. Indicates the second activation boundary; the magnetic latitude activation function satisfy: This represents the local activation function for the magnetic field. When indicating a magnetic place, Indicates the peak value at local magnetic field time. The time span representing the activation of the aurora charging phenomenon; the magnetic local time activation function. satisfy: This represents the difference between the local magnetic time and the peak value. Indicates the local time peak value of the magnetic field.

3. The method as described in claim 1 or 2, characterized in that, The ionization resistance correction model satisfies: in, i Indicates the ionization resistance indicator. This represents the ionization drag experienced by the target aircraft; This represents the ion number density of the target aircraft; Indicates the radius of the target aircraft; Represents a unit charge; Indicates the total charging amount; Indicates the average mass of ions in the environment; This indicates the relative velocity of the target aircraft; Indicates the length of the Debye shield; Indicates the impact parameter; the total charging amount satisfy , This represents the surface charging equilibrium potential of the target aircraft. This indicates the incremental charge of the aurora arc.

4. The method as described in claim 1, characterized in that, The prediction of the orbital decay rate of the target spacecraft in the future time period based on the historical state information and the ionization drag correction model includes: For each of the multiple time points, the two orbital elements, ionization environment parameters, and geomagnetic coordinates of the target spacecraft at that time point are input into the ionization drag correction model to calculate the ionization drag experienced by the target spacecraft at that time point. Calculate the orbital decay rate of the target spacecraft at the time point based on the ionization drag experienced by the target spacecraft at the time point. Integrating the orbital decay rate of the target spacecraft at the multiple time points yields the orbital decay rate of the target spacecraft in the future time period.

5. The method as described in claim 4, characterized in that, The step of calculating the orbital decay rate of the target spacecraft at the time point based on the ionization drag experienced by the target spacecraft at the time point includes: Divide the ionization drag experienced by the target spacecraft at the time point by the mass of the target spacecraft to obtain the ionization drag acceleration of the target spacecraft at the time point. The orbital decay rate of the target spacecraft at the stated time point is calculated using the ionization drag acceleration. The formula for calculating the orbital decay rate is as follows: ,in, The rate of change of the semi-major axis of the orbit. Indicates the orbital decay rate; For the semi-major axis of the track; The gravitational constant of the central celestial body; i Indicates the ionization resistance indicator. This represents the ionization drag acceleration of a target aircraft distributed tangentially.

6. The method as described in claim 1, characterized in that, The acquisition of the historical state information of the target aircraft includes: Obtain the set of two rows of orbital elements corresponding to the available epochs of the target spacecraft within a historical time period; The two sets of orbital elements are densified using a J2 perturbation model to obtain the historical state information of the target spacecraft; the time difference between any two consecutive time points among the multiple time points is less than the time difference between any two consecutive available epochs of the target spacecraft.

7. The method as described in claim 1, characterized in that, The method further includes: The ionization resistance correction model was optimized.

8. A device for predicting the orbital decay rate of a spacecraft based on aurora arc charging enhancement, characterized in that, The device for predicting the orbital decay rate of a target spacecraft affected by ionization drag enhanced by auroral arc charging during a low solar activity year includes an acquisition module and a prediction module. The acquisition module is used to acquire the historical state information of the target spacecraft; the target spacecraft is a low-Earth orbit spacecraft; the historical state information includes the number of two orbital elements of the target spacecraft at each time point within multiple time points of a historical time period; The prediction module is used to predict the orbital decay rate of the target spacecraft in the future time period based on the historical state information and the ionization drag correction model; the ionization drag correction model incorporates the aurora arc charging increment. The ionization resistance model for modifying the aurora arc charging enhancement effect; the aurora arc charging increment It is determined by geomagnetic coordinates and ionization environment parameters; the ionization environment parameters include electron number density, ion number density, and solar activity indicators.

9. An electronic device, characterized in that, The device includes a processor and a memory coupled to the processor; the memory is used to store computer instructions, which, when the electronic device is running, are executed by the processor to cause the electronic device to perform the method as described in any one of claims 1 to 7.

10. A computer-readable storage medium, characterized in that, It includes computer program instructions that, when executed by a computer, cause the computer to perform the method as described in any one of claims 1 to 7.

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