Active Orbit Transfer Method, Equipment and Product for Ultra-Low Orbit Satellite

Abstract

The present application discloses an active orbit transfer method, device and product for ultra-low orbit satellites, relating to the field of satellite detection technology. It includes: when obtaining the first instantaneous orbital elements of the satellite at the first moment and the attitude strategy of the satellite within the adjustment period, inputting the first instantaneous orbital elements and the attitude strategy into the orbital mechanics model to obtain the second instantaneous orbital elements of the satellite at the second moment; converting the second instantaneous orbital elements into mean orbital elements, and taking the difference between the semi-major axis parameter in the mean orbital elements and the target mean orbital semi-major axis to obtain the predicted decay altitude of the satellite within the adjustment period; if the predicted decay altitude is less than zero, start the next adjustment period at the third moment, and the time interval between the first moment and the third moment is T / 2; if the predicted decay altitude is greater than zero and less than or equal to the first decay threshold, control the satellite to raise the predicted decay altitude between the first moment and the third moment, and start the next adjustment period at the third moment.

Description

Technical Field

[0001] The present application relates to the field of space environment detection technology, in particular to the field of satellite detection technology, and in particular to an active orbit change method, equipment and product for an ultra-low orbit satellite. Background Art

[0002] Very low-orbit satellites (LEO satellites) have the advantages of high ground resolution and low signal transmission delay due to their low orbital altitude, and are widely used in earth observation, communication, scientific research and other fields. However, due to the drag of the earth's atmosphere, the LEO satellite orbit will gradually decay, eventually causing the satellite to re-enter the earth's atmosphere and burn up, which will not only cause the loss of expensive satellite equipment, but may also pose a threat to ground facilities and personnel safety.

[0003] At present, the orbit management of most satellites adopts an "after-the-fact response" strategy, that is, the orbit raising operation is carried out after the satellite orbit decays to a certain extent. However, when encountering extreme space weather such as solar flares and geomagnetic explosions, the atmospheric density will change greatly, resulting in a sudden increase in the air resistance of ultra-low orbit satellites. If the satellite is in a relatively low position at this time, it will not be able to be pushed for a long time due to large air resistance, small thrust of ion thrusters, or insufficient power on the satellite, which will cause the satellite to face a greater risk of crashing. Therefore, how to effectively manage the orbit of ultra-low orbit satellites and extend their service life has become a technical problem that needs to be urgently solved in the current field of aerospace technology. Summary of the invention

[0004] In view of this, the embodiments of the present application provide a method, device, equipment, medium and product for active orbit change of ultra-low orbit satellites, which can effectively avoid the risk of satellite crashes and improve the efficiency and reliability of orbit management by predicting the orbit attenuation trend and actively raising the orbit in advance.

[0005] In a first aspect, an embodiment of the present application provides an active orbit transfer method for a very low orbit satellite. The method includes: when obtaining the first instantaneous orbit elements of the satellite at a first moment and the attitude strategy of the satellite within an adjustment period, inputting the first instantaneous orbit elements and the attitude strategy into an orbit mechanics model to obtain the second instantaneous orbit elements of the satellite at a second moment, where the first moment is the start moment of the adjustment period, the second moment is the end moment of the adjustment period, and the period duration of the adjustment period is T; converting the second instantaneous orbit elements into mean orbit elements, and subtracting the semi-major axis parameter in the mean orbit elements from the target mean orbit semi-major axis to obtain the predicted decay altitude of the satellite within the adjustment period, where the target mean orbit semi-major axis is the mean orbit semi-major axis corresponding to the standard altitude of the satellite's operation; if the predicted decay altitude is less than zero, starting the next adjustment period at a third moment, where the time interval between the first moment and the third moment is T / 2; if the predicted decay altitude is greater than zero and less than or equal to a first decay threshold, controlling the satellite to raise the predicted decay altitude between the first moment and the third moment, and starting the next adjustment period at the third moment.

[0006] In a second aspect, an embodiment of the present application provides an active orbit transfer device for a very low orbit satellite. The device includes: an input module, configured to input the first instantaneous orbit elements and the attitude strategy into an orbit mechanics model when obtaining the first instantaneous orbit elements of the satellite at a first moment and the attitude strategy of the satellite within an adjustment period, to obtain the second instantaneous orbit elements of the satellite at a second moment, where the first moment is the start moment of the adjustment period, the second moment is the end moment of the adjustment period, and the period duration of the adjustment period is T; a determination module, configured to convert the second instantaneous orbit elements into mean orbit elements, and subtract the semi-major axis parameter in the mean orbit elements from the target mean orbit semi-major axis to obtain the predicted decay altitude of the satellite within the adjustment period, where the target mean orbit semi-major axis is the mean orbit semi-major axis corresponding to the standard altitude of the satellite's operation; an orbit elevation module, configured to start the next adjustment period at a third moment if the predicted decay altitude is less than zero, where the time interval between the first moment and the third moment is T / 2; the orbit elevation module is further configured to control the satellite to raise the predicted decay altitude between the first moment and the third moment if the predicted decay altitude is greater than zero and less than or equal to a first decay threshold, and start the next adjustment period at the third moment.

[0007] In a third aspect, an embodiment of the present application provides an electronic device, including: a processor and a memory storing computer program instructions; when the processor executes the computer program instructions, the steps of the active orbit transfer method for a very low orbit satellite as in the first aspect are implemented.

[0008] Fourthly, an embodiment of the present application provides a computer-readable storage medium, on which computer program instructions are stored. When the computer program instructions are executed by a processor, the steps of the active orbit change method of the ultra-low orbit satellite as described in the first aspect are implemented.

[0009] Fifthly, an embodiment of the present application provides a computer program product, which is stored in a non-volatile storage medium. When the computer program product is executed by a processor, the steps of the active orbit change method of the ultra-low orbit satellite as described in the first aspect are implemented.

[0010] Sixthly, an embodiment of the present application provides a chip, which includes a processor and a communication interface. The communication interface is coupled to the processor, and the processor is used to run programs or instructions to implement the steps of the active orbit change method of the ultra-low orbit satellite as described in the first aspect.

[0011] The present application provides an active orbit change method, device, equipment, medium and product for an ultra-low orbit satellite. By using the semi-major axis of the mean orbital elements as the judgment index for satellite orbit decay, the decay height of the satellite within a future adjustment period can be predicted in advance to obtain the predicted decay height. When it is predicted that the satellite decays, an orbit raising operation is actively carried out in advance to reduce the falling risk. In this way, by predicting the orbit decay trend and actively raising the orbit in advance, the risk of satellite crash can be effectively avoided, and the efficiency and reliability of orbit management can be improved. BRIEF DESCRIPTION OF THE DRAWINGS

[0012] In order to more clearly illustrate the technical solution of the embodiments of the present application, the following briefly introduces the drawings in the embodiments of the present application.

[0013] Figure 1 It is a schematic flowchart of the active orbit change method of the ultra-low orbit satellite provided by an embodiment of the present application;

[0014] Figure 2 It is a schematic structural diagram of an active orbit change device of the ultra-low orbit satellite provided by an embodiment of the present application;

[0015] Figure 3 It is a schematic hardware structure diagram of an electronic device provided by an embodiment of the present application. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0016] The principles and spirit of the present application will be described below with reference to several exemplary embodiments. It should be understood that the purpose of providing these embodiments is to make the principles and spirit of the present application clearer and more thorough, so that those skilled in the art can better understand and then implement the principles and spirit of the present application. The exemplary embodiments provided herein are only a part of the embodiments of the present application, rather than all the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments herein without creative efforts fall within the scope of protection of the present application.

[0017] In this document, terms such as first, second, and third are only used to distinguish one entity (or operation) from another entity (or operation), and do not require or imply any order or association between these entities (or operations).

[0018] The following briefly describes the relevant content such as the concepts and technical terms that may be involved in the embodiments of the present application.

[0019] Instantaneous orbital elements: are the orbital parameters that strictly satisfy the orbital equation and describe the instantaneous orbital state of a satellite at a specific moment. It includes the influence of all perturbing forces (such as the non-spherical gravity of the Earth, aerodynamic force, and the gravity of the third body, etc.). It describes the actual orbit where the satellite is located at this moment, including the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee, and mean anomaly. Due to the existence of perturbing forces, the instantaneous orbital elements show periodic changes over time.

[0020] Mean orbital elements: are the orbital parameters obtained by eliminating the influence of short-period perturbations and are usually used to describe the orbital characteristics of a satellite on a long-time scale. The mean orbital elements can be derived from the instantaneous orbital elements through a filtering method. It removes the short-period perturbation components and only reflects the long-term and medium-term perturbations. The changes are usually slower and are more suitable for long-term analysis. It describes a smoothed orbit rather than the actual orbit of the satellite at a certain moment.

[0021] In the related art, the orbit management of most satellites adopts a "post-response" strategy, that is, after the satellite orbit decays to a certain extent, an orbit-raising operation is carried out. The specific steps are as follows:

[0022] ① Orbit monitoring: Through the ground TT&C system and on-board sensors, the orbital altitude and decay of the satellite are monitored in real time.

[0023] ② Orbit decay warning: When it is monitored that the orbital altitude of the satellite drops below the preset safety threshold, an orbit decay warning is issued.

[0024] ③ Orbit-raising operation: Start the propulsion system on the satellite to carry out orbit-raising and restore the satellite to the predetermined orbital altitude.

[0025] However, the above-mentioned orbit elevation strategy has the problem of response lag, that is, the orbit elevation operation is carried out after the orbit decays to a certain extent. When encountering extreme space weather such as solar flares and geomagnetic storms, the atmospheric density will change greatly, resulting in a sudden increase in the air resistance suffered by ultra-low orbit satellites. If the satellite is at a lower position at this time (with high resistance and small thrust of the ion thruster), or the on-board power is insufficient (unable to push for a long time), the satellite will face a greater risk of crashing.

[0026] Based on this, by predicting the orbit decay trend and actively elevating the orbit in advance, this application can effectively avoid the risk of satellite crashing, improve the efficiency and reliability of orbit management, and solve the above technical problems.

[0027] The following combines the drawings and details the active orbit transfer method of the ultra-low orbit satellite provided by the embodiments of this application through specific embodiments and their application scenarios.

[0028] Figure 1 It is a schematic flowchart of the active orbit transfer method of the ultra-low orbit satellite provided by an embodiment of this application. The execution subject of the active orbit transfer method of the ultra-low orbit satellite can be an orbit control system in the ultra-low orbit guarantee system.

[0029] The following takes the execution subject of the active orbit transfer method of the ultra-low orbit satellite as the orbit control system as an example to illustrate the active orbit transfer method of the ultra-low orbit satellite of this application. It should be noted that the above execution subject and application scenario do not limit this application.

[0030] As Figure 1 shown, the active orbit transfer method of the ultra-low orbit satellite provided by the embodiments of this application may include Step 110 - Step 140.

[0031] Step 110, when obtaining the first instantaneous orbit elements of the satellite at the first moment and the attitude strategy of the satellite during the adjustment period, input the first instantaneous orbit elements and the attitude strategy into the orbit mechanics model to obtain the second instantaneous orbit elements of the satellite at the second moment;

[0032] Step 120, convert the second instantaneous orbit elements into mean orbit elements, and subtract the semi-major axis parameter in the mean orbit elements from the target mean orbit semi-major axis to obtain the predicted decay height of the satellite during the adjustment period;

[0033] Step 130, if the predicted decay height is less than zero, start the next adjustment period at the third moment;

[0034] Step 140, if the predicted decay height is greater than zero and less than or equal to the first decay threshold, control the satellite to lift the predicted decay height between the first moment and the third moment, and start the next adjustment period at the third moment.

[0035] The active orbit transfer method of the ultra-low orbit satellite provided by the embodiment of the present application uses the semi-major axis of the mean orbital elements as the judgment index for satellite orbit decay, can predict in advance the decay height of the satellite within a future adjustment period, obtain the predicted decay height, and when the satellite decay is predicted, actively perform an orbit raising operation in advance to reduce the falling risk. In this way, by predicting the orbit decay trend and actively raising the orbit in advance, the risk of satellite crash can be effectively avoided, and the efficiency and reliability of orbit management can be improved.

[0036] The following combines specific embodiments to introduce in detail the specific implementation manners of the above steps.

[0037] Regarding step 110, when the first instantaneous orbital elements of the satellite at the first moment and the attitude strategy of the satellite within the adjustment period are obtained, the first instantaneous orbital elements and the attitude strategy are input into the orbit mechanics model to obtain the second instantaneous orbital elements of the satellite at the second moment.

[0038] In step 110, the first instantaneous orbital elements are the satellite orbital elements of the satellite at the first moment, and the attitude strategy of the satellite within the adjustment period is the planned attitude parameters pre-input by relevant personnel. This attitude strategy is used to control the running attitude of the satellite within the adjustment period and includes the pitch angle, roll angle, and yaw angle.

[0039] The first moment is the start moment of the adjustment period, and the second moment is the end moment of the adjustment period. The first moment is the current moment, and the second moment is the future moment. The second instantaneous orbital elements are the satellite orbital elements of the satellite at the second moment, which are predicted values calculated by the orbit mechanics model in combination with the first instantaneous orbital elements and the attitude strategy.

[0040] Satellite orbital elements are a set of six key parameters required to describe the motion of celestial bodies or spacecraft on their Kepler orbits, including the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee, and true anomaly. These parameters can accurately define the orbital position and velocity of an object in three-dimensional space.

[0041] The orbit mechanics model can be a Simplified General Perturbation Model 4 (SGP4), a High Precision Orbit Propagator (HPOP), etc. The present application does not make specific limitations on this.

[0042] It should be noted that the prediction period of the orbital mechanics model is shorter than the adjustment period. The orbital mechanics model does not directly predict the second instantaneous orbital elements based on the first instantaneous orbital elements, but multiple predictions are required to obtain them. For example, if the prediction period is 3h and the adjustment period is 24h, then the first prediction obtains the instantaneous orbital elements at the 3rd hour, the second prediction obtains the instantaneous orbital elements at the 6th hour (predicted values). After multiple predictions, the instantaneous orbital elements at the 24th hour, that is, the second instantaneous orbital elements, are obtained based on the instantaneous orbital elements at the 21st hour (predicted values).

[0043] The period of the adjustment period is T, and this adjustment period is used to predict the decay altitude of the satellite and adjust the real-time operating altitude of the satellite based on this decay altitude.

[0044] Since the lower the altitude of the satellite, the greater the risk of destruction under the passive orbit-raising strategy, the period is set to be positively correlated with the working altitude (or operating altitude) of the satellite. That is, the lower the altitude of the satellite, the shorter the adjustment period. That is, the real-time operating altitude of the satellite is adjusted at a higher adjustment frequency, and when the satellite decay is predicted, the orbit-raising operation can be carried out in advance in a timely manner to reduce the falling risk.

[0045] Exemplarily, when the satellite operates at an altitude of 300km, the recommended value of the adjustment period T is 24 hours. As the working altitude of the satellite decreases, the value of T needs to become smaller.

[0046] Involved in step 120, the second instantaneous orbital elements are converted into mean orbital elements, and the difference between the semi-major axis parameter in the mean orbital elements and the target mean orbital semi-major axis is obtained to get the predicted decay altitude of the satellite within the adjustment period.

[0047] In step 120, the semi-major axis parameter in the second instantaneous orbital elements is the instantaneous orbital semi-major axis, and the semi-major axis parameter in the mean orbital elements is the instantaneous mean orbital semi-major axis. In this application, the Fourier transform algorithm or the time smoothing algorithm can be used to perform time averaging on the instantaneous orbital semi-major axis in the second instantaneous orbital elements to obtain the instantaneous mean orbital semi-major axis corresponding to the satellite at the second moment. The predicted decay altitude is the predicted value of the decay altitude of the satellite after one adjustment period.

[0048] Exemplarily, at the start time t0 of the adjustment period, the first instantaneous orbital elements of the satellite are obtained through telemetry data. Taking the attitude strategy of the satellite in this period as the input, the orbit of the satellite is extrapolated for a period of T to obtain the second instantaneous orbital elements of the satellite at the moment t0+T, and the result is converted to obtain the instantaneous mean orbital semi-major axis , then the predicted decay altitude , is the target mean orbital semi-major axis.

[0049] Converting the semi-major axis of the instantaneous orbital elements to that of the mean orbital elements is a complex process because the instantaneous orbital elements (such as position, velocity, etc.) vary with time, while the mean orbital elements (such as mean semi-major axis, mean eccentricity, etc.) are obtained by averaging over time. This conversion usually involves knowledge of orbital mechanics and perturbation theory. In orbital mechanics, the instantaneous orbital elements typically include six parameters: semi-major axis, eccentricity, inclination, right ascension of the ascending node, argument of perigee, and true anomaly. These parameters describe the orbital state of a satellite at a certain moment. However, due to factors such as the non-sphericity of the Earth, atmospheric drag, and gravitational perturbations from the Sun and the Moon, the orbit of the satellite changes over time. Therefore, the instantaneous orbital elements also change with time. To obtain the mean orbital elements, it is necessary to perform a time average on the instantaneous orbital elements. This usually involves numerically integrating the satellite orbit and calculating the average value of each orbital element over a certain time interval.

[0050] It should be noted that when calculating the predicted decay altitude in this application, only the semi-major axis parameter in the mean orbital elements is used. Therefore, it is only necessary to convert the semi-major axis parameter in the second instantaneous orbital elements to the instantaneous mean semi-major axis, and there is no need to convert the other five parameters.

[0051] In some embodiments of this application, the period duration of the oscillation period is T0. Using Fourier transform to perform a time average on the instantaneous orbital semi-major axis in the second instantaneous orbital elements to obtain the instantaneous mean orbital semi-major axis corresponding to the satellite at the second moment may specifically include: obtaining all the instantaneous orbital elements of the satellite within at least one oscillation period before the second moment to obtain a data set; performing Fourier transform on the second instantaneous orbital elements and all the instantaneous orbital semi-major axes in the data set to obtain the frequency-domain representation of the orbital semi-major axis; filtering the frequency-domain representation based on a low-pass filter to truncate the high-frequency components, only retaining the DC component and low-frequency components (representing the average value and long-term changes), and filtering out the high-frequency periodic oscillations; performing inverse Fourier transform on the filtered frequency-domain representation to obtain the filtered time series; using the filtered time series as the instantaneous mean orbital semi-major axis of the satellite at the second moment.

[0052] In some other embodiments of this application, the period duration of the oscillation period is T0. Using a time smoothing algorithm to perform a time average on the instantaneous orbital semi-major axis in the second instantaneous orbital elements to obtain the instantaneous mean orbital semi-major axis corresponding to the satellite at the second moment may specifically include: taking the second moment as the middle moment and the period duration T0 as the time window length to determine the target time window; obtaining all the instantaneous orbital elements of the satellite within the target time window and performing a moving average calculation on the average value of the instantaneous orbital semi-major axes among all the instantaneous orbital elements within the target time window to obtain the instantaneous mean orbital semi-major axis of the satellite at the second moment.

[0053] It should be noted that when obtaining all the instantaneous orbital elements of the satellite within the target time window as described above, the orbital mechanics model also needs to predict the instantaneous orbital elements within the time period of T0 / 2 after the second moment.

[0054] The average semi-major axis of the satellite's orbit is an important parameter, which determines the orbital shape and period of the satellite orbiting the Earth. The target average semi-major axis is the average semi-major axis corresponding to the standard altitude h of the satellite's operation (i.e., the average altitude of the satellite from the Earth's surface), and the standard altitude of the satellite's operation is preset. ≈r = R + h, where R is the average radius of the Earth (about 6371 km) and r is the orbital radius of the satellite.

[0055] Exemplarily, if h is the average altitude of 300 km, then = 6371 + 300 = 6671 km.

[0056] Regarding step 130, if the predicted decay height is less than zero, then the next adjustment cycle is started at the third moment.

[0057] In step 130, the time interval between the first moment and the third moment is T / 2. If the predicted decay height is less than zero, it indicates that the satellite has no risk of falling temporarily within the next adjustment cycle. Therefore, no orbit-raising operation is performed on the satellite, and the next adjustment cycle is started in advance by T / 2. At the start moment of the next adjustment cycle, the future decay height of the satellite is predicted, and whether to perform an orbit-raising operation is judged based on this future decay height.

[0058] Exemplarily, if T is 24 h and the first moment is 0:01 on February 1st, if the predicted decay height of the satellite from 0:01 to 24:00 on February 1st is less than zero, then the next adjustment cycle can be started at 12:01 on February 1st, and the decay height of the satellite in the next 24 h is predicted at 12:01 on February 1st.

[0059] Regarding step 140, if the predicted decay height is greater than zero and less than or equal to the first decay threshold, then the satellite is controlled to raise the predicted decay height between the first moment and the third moment, and the next adjustment cycle is started at the third moment.

[0060] In step 140, the first decay threshold can be calculated in advance or set in advance, and the present application does not make specific limitations on this.

[0061] It should be noted that the above step 130 and step 140 are two judgment branches, so the present application does not make specific limitations on their execution order.

[0062] In some embodiments of the present application, the first attenuation threshold can be pre-calculated. Before obtaining the predicted attenuation height of the satellite within the adjustment period in step 120 above, it may further include:

[0063] Obtain the average attenuation speed of the satellite within the previous adjustment period before the first moment, and determine the average attenuation speed within the previous adjustment period as the first attenuation threshold.

[0064] In the embodiments of the present application, the satellite orbit attenuation speed is affected by the solar activity cycle, which is approximately 11 years and is divided into solar activity high years and low years. During solar activity high years, solar eruptions are frequent, the atmospheric density increases, and the satellite orbit attenuation speed significantly accelerates. Therefore, the present application can determine the average attenuation speed of the satellite within the previous adjustment period before the first moment as the first attenuation threshold. Thus, the first attenuation threshold can vary with solar activity. For example, for a satellite at an altitude of 300 km, the attenuation speed is about 1 - 2 km / day during solar quiet periods and can reach 5 km / day during solar eruption periods, enabling the magnitude of the first attenuation threshold to better conform to the operating state of the satellite under the current solar activity state.

[0065] In some embodiments of the present application, in order to accurately lift the satellite to a specified orbital height, when performing an orbit-raising operation on the satellite, it may include: calculating the target duration required for the satellite to be lifted to the specified height through the circular orbit formula, and splitting the target duration into multiple time periods to control the engine to start thrusting within the multiple time periods to lift the satellite.

[0066] Specifically, according to the formula of the circular orbit, it can be known that the tangential velocity of the satellite in the orbit , then to lift a satellite on a circular orbit with a semi-major axis of to a circular orbit with a semi-major axis of , the change in its linear velocity . The target duration of starting thrusting can be further calculated based on the engine thrust F of the satellite and the satellite mass m: . After that, it can be divided into multiple time periods for starting thrusting, and the sum of the durations is t.

[0067] In the embodiments of the present application, through the circular orbit formula, the target duration required for the satellite to be lifted to the specified height can be accurately calculated. Based on this, splitting the target duration into multiple time periods and controlling the engine to start thrusting within the multiple time periods can enable the satellite to be accurately lifted to the specified orbital height.

[0068] In some other embodiments of the present application, in order to accurately lift a satellite to a specified height, when performing an orbit-raising operation on the satellite, it may include: during the continuous thrust of the engine, converting the instantaneous orbital elements fed back by the satellite in real time into mean orbital elements, determining the actual lifting height of the satellite based on the semi-major axis parameter in the mean orbital elements, and continuing to thrust when the actual lifting height is less than the specified height until the actual lifting height reaches the specified height.

[0069] In the embodiments of the present application, the engine can be controlled to continuously thrust. During the thrust, the actual lifting height of the satellite can be accurately determined based on the instantaneous orbital elements fed back by the satellite in real time, so as to ensure continuous thrust when the actual lifting height does not reach the specified height, thereby realizing lifting the satellite to the specified height at one time without stopping in the middle, with a faster orbit-raising speed and higher efficiency.

[0070] In some embodiments of the present application, in order to timely remedy the satellite when the attenuation degree of the satellite is relatively high, after obtaining the predicted attenuation height of the satellite within the adjustment period in step 120 above, it may further include:

[0071] Step 150, if the predicted attenuation height is greater than the first attenuation threshold and less than or equal to the second attenuation threshold, then control the satellite to lift to the target height between the first moment and the third moment, and start the next adjustment period at the third moment.

[0072] Wherein, the target height satisfies the following conditions: greater than or equal to the first attenuation threshold and less than or equal to the second attenuation threshold and less than the first height, the first height being the lifting height of the satellite after continuous thrust within T / 4 obtained based on the circular orbit formula and the second attenuation threshold being twice the first attenuation threshold.

[0073] In the embodiments of the present application, the first attenuation threshold is the average attenuation speed of the satellite within the previous adjustment period before the first moment. Therefore, the first attenuation threshold is set as the minimum value of the target height, so that the lifting height of the satellite matches the attenuation speed of the satellite within the adjustment period. If the actual attenuation height does not reach the predicted value, the satellite will deviate from its original orbit if it is lifted too high (at an orbital altitude of 300 km, a height difference of 2 km will cause the deviation magnitude of the satellite's position to reach hundreds of kilometers per day). Therefore, by limiting the target height to be less than or equal to twice the first attenuation threshold, the lifting height of the satellite is restricted to prevent the satellite from being lifted too high. Considering the lifting ability design of the satellite, it is less than the height lifted by the satellite during continuous thrust within a quarter of the adjustment period, so that the lifting height of the satellite conforms to the engine's ability and prevents the target height from being set too high beyond the engine's operating ability.

[0074] Thus, a strategy of small - amplitude lift and high - frequency prediction is adopted. High - frequency means that the period of the lift operation is shorter than the entire adjustment period T. The frequency decreases as the decay risk increases (T / 2, T / 4 or smaller). The advantage of this is that when the orbital decay prediction is too small or too large, the next adjustment cycle can be started in time, thus reducing the risk brought by the prediction error. Small - amplitude lift means that the height of each lift is not higher than twice the first decay threshold. This avoids the orbital deviation caused by the satellite being lifted too high when the predicted decay amount is too large compared to the actual decay amount. When the actual decay risk is too high, the high - frequency compensates for the deficiency of the lift amplitude. In summary, this strategy for setting the target height increases the robustness of the overall orbital lift scheme and weakens the risk brought by the orbital prediction error.

[0075] In the related art, when encountering extreme space weather such as solar flares and geomagnetic storms, the atmospheric density will change greatly, resulting in a sudden increase in the air resistance suffered by ultra - low - orbit satellites. If the satellite is at a relatively low position at this time, due to the large air resistance and the small thrust of the ion thruster, it is impossible to keep the thruster on for a long time, and thus the satellite faces a greater risk of crashing.

[0076] In some embodiments of the present application, before controlling the satellite to lift the target height between the first moment and the third moment in step 150 above, it may further include: obtaining the solar activity phase at the first moment; in the case where the solar activity phase is a quiet period, determining the first decay threshold as the target height; in the case where the solar activity phase is a peak period, taking the minimum value of the second decay threshold and the first height as the maximum lift, and determining the maximum lift as the target height; in the case where the solar activity phase is a descending period, determining the sum value of the first decay threshold and the lift adjustment amount as the target height; in the case where the solar activity phase is an ascending period, determining the sum value of the first decay threshold and twice the lift adjustment amount as the target height.

[0077] Among them, the lift adjustment amount is the difference between the first decay threshold and the maximum lift divided by three.

[0078] In the embodiments of the present application, the height of the target altitude is related to the intensity of solar activity. The more intense the solar activity is, the higher the target altitude is. Thus, when solar activity is relatively calm, the air resistance on the ultra-low orbit satellite is relatively small, the crash risk of the satellite is relatively low, and the safety factor is relatively high. At this time, within a reasonable range, a relatively low lift altitude can be set for the satellite to avoid the satellite getting out of its original orbit due to excessive lift. When solar activity is relatively intense, the air resistance on the ultra-low orbit satellite is relatively large, the crash risk of the satellite is relatively high, and the safety factor is relatively low. At this time, within a reasonable range, a relatively high lift altitude can be set for the satellite to avoid the situation where the satellite cannot maintain thrust for a long time due to large air resistance and small ion thruster thrust when it is at a low position, and reduce the crash risk of the satellite.

[0079] In some embodiments of the present application, after obtaining the predicted decay altitude of the satellite within the adjustment period in step 120 above, it may further include: if the predicted decay altitude is greater than the second decay threshold and less than or equal to the third decay threshold, then control the satellite to lift the target altitude between the first moment and the fourth moment, and start the next adjustment period at the fourth moment, where the third decay threshold is the average decay rate of the satellite during the adjustment period duration in the solar burst period, the third decay threshold is not less than four times the first decay threshold, and the interval duration between the first moment and the fourth moment is T / 4.

[0080] In the embodiments of the present application, the third decay threshold is the average decay rate of the satellite during the adjustment period duration in the solar burst period. Therefore, this third decay threshold represents a relatively high decay degree. If the predicted decay altitude of the satellite is between the second decay threshold and the third decay threshold, it indicates that the decay degree of the satellite within the adjustment period is relatively serious. Based on this, the present application further shortens the orbit lift time of the satellite and accelerates the orbit lift process of the satellite, and quickly lifts the target altitude of the satellite within T / 4 after the first moment. And, start the next adjustment period after T / 4, making the period of the lift operation shorter relative to the entire adjustment period T. When the orbit decay prediction is too large, the next adjustment period can be started in time, thereby reducing the risk brought by the prediction error.

[0081] In other embodiments of the present application, before obtaining the predicted decay altitude of the satellite within the adjustment period in step 120 above, it may further include: obtaining the average decay rates of the satellite in multiple adjustment periods before the first moment when the solar activity phase is at the peak period, and determining the maximum value of the multiple average decay rates as the third decay threshold.

[0082] In some embodiments of the present application, after obtaining the predicted decay height of the satellite within the adjustment period in step 120 above, it may further include: if the predicted decay height is greater than the third decay threshold, determining a first value n based on the predicted decay height and the target height; controlling the satellite to raise the target height between the first moment and the fifth moment, and starting the next adjustment period at the fifth moment.

[0083] Specifically, , Taking [target height] as the target height, the time interval between the first moment and the fifth moment is T / n.

[0084] In the embodiments of the present application, if the predicted decay height of the satellite is greater than the third decay threshold, it indicates that the decay degree of the satellite within the adjustment period is serious. Therefore, it is necessary to further shorten the time spent on raising the satellite's orbit. By taking the integer part of the predicted decay height divided by the target height and adding 1, within T / n after the first moment, the satellite is quickly raised to the target height, further accelerating the process of raising the satellite's orbit. Moreover, the next adjustment period is started after T / n, making the period of the raising operation shorter relative to the entire adjustment period T. When the orbit decay prediction is too large, the next adjustment period can be started in a timely manner, thereby reducing the risk brought by the prediction error.

[0085] As a specific example, T = 24h, the predicted decay height of the satellite within the adjustment period is , and the first decay threshold, the second decay threshold, and the third decay threshold are respectively , and the target height is , then the orbit raising strategy provided by the present application is as follows:

[0086] ① , indicating that the algorithm predicts that the height of the satellite after one cycle is above the standard height. At this time, the satellite does not need to perform a raising operation, and the next adjustment period is started at the moment;

[0087] ② , indicating that the predicted decay height is slightly lower than the standard height. At this time, the satellite needs to be raised by within 12 hours, and the next adjustment period is started at the moment;

[0088] ③ , at this time the satellite needs to be raised by within 12 hours, and the next adjustment period is started at the moment;

[0089] ④ , at this time the satellite needs to be raised by within 6 hours, and the next adjustment period is started at the moment;

[0090] ⑤ It indicates that the predicted decay height shows a large amplitude of decay in the orbit within one cycle. At this time, human intervention can be added. If there is no human intervention, it is necessary to further increase the adjustment frequency. , The satellite needs to lift within and start the next adjustment cycle at the

[0091] In some embodiments of the present application, the accurate calculation of the predicted decay height depends on the accurate prediction of the second instantaneous orbital elements. Based on this, in order to improve the prediction accuracy of the second instantaneous orbital elements, the present application can correct the second instantaneous orbital elements output by the orbital mechanics model based on the orbit correction model. The specific training steps of the above orbit correction model are as follows:

[0092] Obtain the time characteristics, satellite state characteristics, space environment characteristics, and satellite orbital elements at multiple first moments; input the satellite orbital elements at the first moment into the orbital mechanics model to obtain the predicted values of the velocity position parameters for the second moment output by the orbital mechanics model, where the time interval between the first moment and the second moment is a preset cycle duration; combine the time characteristics, satellite state characteristics, and space environment characteristics at each first moment, as well as the predicted values and true values of the velocity position parameters for the second moment, to construct historical training data; use the time characteristics, satellite state characteristics, and space environment characteristics at the first moment, as well as the predicted values of the velocity position parameters for the second moment as input training data, and use the difference between the predicted values and true values of the velocity position parameters for the second moment as output training data (i.e., output labels) for training.

[0093] Specifically, the first moment and the second moment are two adjacent historical prediction moments, and the second moment is after the first moment. The true values of the velocity position parameters for the second moment can be obtained through actual detection of the satellite by the satellite detection system. When training the model, the mean square error can be used as the loss function, and the backpropagation algorithm can be used for training. The orbit correction model is associated with the orbital mechanics model. The preset model can be set according to specific requirements, and the present application does not make specific limitations on this. The difference between the predicted values and true values of the velocity position parameters for the second moment can be the true error value. Calculate the mean square error between the training error prediction value output by the preset model and the true error value to obtain the loss function value.

[0094] It should be noted that in the historical training data, the satellite state features are in the ECEF coordinate system, and the space environment features are in the geodetic coordinate system. The predicted values and true values of the velocity and position parameters at the second moment are both in the ECEF coordinate system, ensuring that the historical training data are uniformly represented in the geocentric coordinate system.

[0095] In the embodiment of the present application, during the training process of the orbit correction model, the training data used is from the predicted values actually output by the orbit mechanics model. Based on this, during the training process of the orbit correction model, it can learn the deviations between the predicted values and true values output by the orbit mechanics model under different time features, different satellite state features, and different space environment features. Furthermore, the trained orbit correction model can have the ability to reasonably and accurately evaluate the errors of the predicted values output by the orbit mechanics model in different satellite orbit prediction scenarios by combining time features, satellite state features, and space environment features, and obtain an orbit correction model adapted to the orbit mechanics model, enabling the orbit correction model to have a better prediction error evaluation ability for the associated orbit mechanics model.

[0096] Corresponding to the method embodiment of the present application, the present application also provides an active orbit transfer device for a very low orbit satellite.

[0097] Figure 2 It is a schematic structural diagram of an active orbit transfer device for a very low orbit satellite provided by an embodiment of the present application. As Figure 2 shown, the active orbit transfer device 200 of the very low orbit satellite may include: an input module 210, a determination module 220, and an orbit elevation module 230.

[0098] Among them, the input module 210 is used to input the first instantaneous orbital elements of the satellite at the first moment and the attitude strategy during the adjustment period to the orbit mechanics model when the first instantaneous orbital elements of the satellite at the first moment and the attitude strategy during the adjustment period are obtained, and obtain the second instantaneous orbital elements of the satellite at the second moment, where the first moment is the start moment of the adjustment period, the second moment is the end moment of the adjustment period, and the period duration of the adjustment period is T; the determination module 220 is used to convert the second instantaneous orbital elements into average orbital elements, and calculate the difference between the semi-major axis parameter in the average orbital elements and the target average orbital semi-major axis to obtain the predicted decay height of the satellite during the adjustment period, and the target average orbital semi-major axis is the average orbital semi-major axis corresponding to the standard operating height of the satellite; the orbit elevation module is used to, if the predicted decay height is less than zero, start the next adjustment period at the third moment, where the time interval between the first moment and the third moment is T / 2; the orbit elevation module 230 is further used to, if the predicted decay height is greater than zero and less than or equal to the first decay threshold, control the satellite to elevate the predicted decay height between the first moment and the third moment, and start the next adjustment period at the third moment.

[0099] The active orbit transfer device for ultra-low orbit satellites provided by the embodiments of the present application uses the semi-major axis of the mean orbital elements as the judgment index for satellite orbit decay, can predict in advance the decay height of the satellite within a future adjustment period, obtain the predicted decay height, and when it is predicted that the satellite decays, actively perform an orbit raising operation in advance to reduce the falling risk. In this way, by predicting the orbit decay trend and actively raising the orbit in advance, the risk of satellite crash can be effectively avoided, and the efficiency and reliability of orbit management can be improved.

[0100] In some embodiments of the present application, the orbit raising module is further configured to, after obtaining the predicted decay height of the satellite within the adjustment period, if the predicted decay height is greater than the first decay threshold and less than or equal to the second decay threshold, control the satellite to raise the target height between the first moment and the third moment, and start the next adjustment period at the third moment; wherein, the target height satisfies the following conditions: greater than or equal to the first decay threshold, less than or equal to the second decay threshold, and less than the first height, and the first height is the raised height of the satellite after continuously pushing for T / 4 obtained based on the circular orbit formula, and the second decay threshold is twice the first decay threshold.

[0101] In some embodiments of the present application, the orbit raising module is further configured to: obtain the solar activity phase at the first moment; in the case that the solar activity phase is a quiet period, determine the first decay threshold as the target height; in the case that the solar activity phase is a peak period, take the minimum value of the second decay threshold and the first height as the maximum raising value, and determine the maximum raising value as the target height; in the case that the solar activity phase is a descending period, determine the sum value of the first decay threshold and the raising adjustment amount as the target height; in the case that the solar activity phase is an ascending period, determine the sum value of the first decay threshold and twice the raising adjustment amount as the target height; wherein, the raising adjustment amount is the difference between the first decay threshold and the maximum raising value divided by three.

[0102] In some embodiments of the present application, the orbit raising module is further configured to: after obtaining the predicted decay height of the satellite within the adjustment period, if the predicted decay height is greater than the second decay threshold and less than or equal to the third decay threshold, control the satellite to raise the target height between the first moment and the fourth moment, and start the next adjustment period at the fourth moment, wherein the third decay threshold is the average decay speed of the satellite during the adjustment period duration in the solar burst period, the third decay threshold is not less than four times the first decay threshold, and the interval duration between the first moment and the fourth moment is T / 4.

[0103] In some embodiments of the present application, the determination module is further configured to, before obtaining the predicted decay altitude of the satellite within the adjustment period, when the solar activity phase is at the peak period, obtain multiple average decay rates of the satellite within multiple adjustment periods before the first moment, and determine the maximum value among the multiple average decay rates as the third decay threshold.

[0104] In some embodiments of the present application, the orbit elevation module is further configured to: after obtaining the predicted decay altitude of the satellite within the adjustment period, if the predicted decay altitude is greater than the third decay threshold, determine a first value n based on the predicted decay altitude and the target altitude; control the satellite to elevate the target altitude between the first moment and the fifth moment, and start the next adjustment period at the fifth moment, where the time interval between the first moment and the fifth moment is T / n.

[0105] In some embodiments of the present application, the orbit elevation module is further configured to: calculate the target duration required for the satellite to elevate a specified altitude through the circular orbit formula, and split the target duration into multiple time periods, so as to control the engine to start thrusting within the multiple time periods to elevate the satellite.

[0106] In some embodiments of the present application, the determination module is further configured to, before obtaining the predicted decay altitude of the satellite within the adjustment period, obtain the average decay rate of the satellite within the previous adjustment period before the first moment, and determine the average decay rate within the previous adjustment period as the first decay threshold.

[0107] The active orbit transfer device for ultra-low orbit satellites provided by the embodiments of the present application can achieve Figure 1 each process implemented by the service platform in the method embodiments, and can achieve the same technical effects. To avoid repetition, it will not be elaborated here.

[0108] Figure 3 It is a schematic hardware structure diagram of an electronic device provided by an embodiment of the present application.

[0109] As Figure 3 shown, the electronic device 300 includes a memory 301, a processor 302, and a computer program stored on the memory 301 and executable on the processor 302.

[0110] In one example, the above-mentioned processor 302 may include a central processing unit (CPU), or an application specific integrated circuit (ASIC), or one or more integrated circuits configured to implement the embodiments of the present application.

[0111] The memory 301 may include a read-only memory (ROM), a random access memory (RAM), a magnetic disk storage medium device, an optical storage medium device, a flash memory device, an electrical, optical, or other physical / tangible memory storage device. Thus, generally, the memory includes one or more tangible (non-transitory) computer-readable storage media (e.g., memory devices) encoded with software including computer-executable instructions, and when the software is executed (e.g., by one or more processors), it is operable to perform the operations described with reference to the method in the embodiment of the first aspect of the present application.

[0112] The processor 302 runs a computer program corresponding to the executable program code by reading the executable program code stored in the memory 301 to implement the method in the embodiment of the first aspect described above.

[0113] In some examples, the electronic device 300 may further include a communication interface 303 and a bus 310. Among them, as Figure 3 shown, the memory 301, the processor 302, and the communication interface 303 are connected through the bus 310 and complete communication with each other.

[0114] The communication interface 303 is mainly used to implement communication between the various modules, devices, units, and / or devices in the embodiments of the present application. The input device and / or output device may also be accessed through the communication interface 303.

[0115] The bus 310 includes hardware, software, or both, and couples the components of the electronic device 300 to each other. By way of example and not limitation, the bus 310 can include an Accelerated Graphics Port (AGP) or other graphics bus, an Enhanced Industry Standard Architecture (EISA) bus, a Front Side Bus (FSB), a Hyper Transport (HT) interconnect, an Industry Standard Architecture (ISA) bus, an InfiniBand interconnect, a Low pin count (LPC) bus, a memory bus, a Micro Channel Architecture (MCA) bus, a Peripheral Component Interconnect (PCI) bus, a PCI-Express (PCI-E) bus, a Serial Advanced Technology Attachment (SATA) bus, a Video Electronics Standards Association Local Bus (VLB) bus, or other suitable buses or a combination of two or more of these. Where appropriate, the bus 310 can include one or more buses. Although embodiments of the present application describe and illustrate specific buses, the present application contemplates any suitable bus or interconnect.

[0116] The electronic device provided by the embodiments of the present application can implement Figure 1 each process implemented by the electronic device in the method embodiment, and can achieve the same technical effects. To avoid repetition, it will not be described in detail here.

[0117] In combination with the active orbit change method of the ultra-low orbit satellite in the above embodiments, the embodiments of the present application can provide a computer storage medium to implement. Computer program instructions are stored on the computer storage medium; when the computer program instructions are executed by a processor, the steps of any one of the active orbit change methods of the ultra-low orbit satellite in the above embodiments are implemented.

[0118] In combination with the active orbit change method of the ultra-low orbit satellite in the above embodiments, the embodiments of the present application can provide a computer program product to implement. The (computer) program product is stored in a non-volatile storage medium, and when the program product is executed by at least one processor, the steps of any one of the active orbit change methods of the ultra-low orbit satellite in the above embodiments are implemented.

[0119] Another embodiment of the present application provides a chip, which includes a processor and a communication interface. The communication interface is coupled to the processor. The processor is used to run programs or instructions to implement each process of the above-mentioned embodiment of the active orbit change method for ultra-low orbit satellites and can achieve the same technical effects. To avoid repetition, it will not be elaborated here.

[0120] It should be understood that the chip mentioned in the embodiments of the present application may also be referred to as a system-on-chip, system chip, chip system, or system-on-chip, etc.

[0121] It should be clear that the present application is not limited to the specific configurations and processes described above and shown in the figures. For the sake of brevity, the detailed description of known methods is omitted here. In the above embodiments, several specific steps are described and shown as examples. However, the method process of the present application is not limited to the specific steps described and shown. Those skilled in the art can make various changes, modifications, and additions, or change the order of the steps after understanding the spirit of the present application.

[0122] It should also be noted that the functional blocks shown in the above structural block diagrams can be implemented as hardware, software, firmware, or a combination thereof. When implemented in hardware, it can be, for example, an electronic circuit, an application specific integrated circuit (ASIC), appropriate firmware, a plug-in, a functional card, and so on. When implemented in software, the elements of the present application are programs or code segments used to perform the required tasks. The program or code segment can be stored in a machine-readable medium or transmitted via a data signal carried in a carrier wave on a transmission medium or a communication link. "Machine-readable medium" can include any medium that can store or transmit information. Examples of machine-readable media include electronic circuits, semiconductor memory devices, ROM, flash memory, erasable ROM (EROM), floppy disks, CD-ROMs, optical discs, hard disks, fiber optic media, radio frequency (RF) links, and so on. The code segment can be downloaded via a computer network such as the Internet, intranet, etc.

[0123] It also needs to be noted that the exemplary embodiments mentioned in the present application describe some methods or systems based on a series of steps or devices. However, the present application is not limited to the order of the above steps, that is, the steps can be executed in the order mentioned in the embodiments, or different from the order in the embodiments, or several steps can be executed simultaneously.

[0124] Aspects of the present disclosure have been described above with reference to the flowcharts and / or block diagrams of methods, apparatuses (systems), and computer program products according to embodiments of the present disclosure. It should be understood that each block in the flowchart and / or block diagram, and the combinations of blocks in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, or other programmable data processing apparatus to produce a machine, such that the instructions executed by the processor of the computer or other programmable data processing apparatus enable the implementation of the functions / actions specified in one or more blocks of the flowchart and / or block diagram. Such a processor can be, but is not limited to, a general-purpose processor, a special-purpose processor, a special application processor, or a field programmable logic circuit. It should also be understood that each block in the block diagram and / or flowchart, and the combinations of blocks in the block diagram and / or flowchart, can also be implemented by dedicated hardware that performs the specified functions or actions, or by a combination of dedicated hardware and computer instructions.

[0125] As described above, the above is only a specific implementation manner of the present application. Those skilled in the art can clearly understand that for the convenience and brevity of description, the specific working processes of the systems, modules, and units described above can refer to the corresponding processes in the foregoing method embodiments, and will not be described herein again. It should be understood that the protection scope of the present application is not limited thereto. Any person skilled in the art can easily think of various equivalent modifications or substitutions within the technical scope disclosed by the present application, and these modifications or substitutions should all be covered by the protection scope of the present application.

Claims

1. An active orbit change method for an ultra-low orbit satellite, characterized in that: include: When a first instantaneous orbital element of the satellite at a first moment and an attitude strategy of the satellite in an adjustment period are obtained, the first instantaneous orbital element and the attitude strategy are input into an orbital mechanics model to obtain a second instantaneous orbital element of the satellite at a second moment, wherein the first moment is a start moment of the adjustment period, the second moment is an end moment of the adjustment period, and the period length of the adjustment period is T; Converting the second instantaneous orbital element into an average orbital element, and subtracting the semi-major axis parameter in the average orbital element from the target average orbital semi-major axis to obtain a predicted attenuation height of the satellite within the adjustment period, wherein the target average orbital semi-major axis is the average orbital semi-major axis corresponding to the standard altitude of the satellite; If the predicted attenuation height is less than zero, the next adjustment cycle is started at the third moment, wherein the interval between the first moment and the third moment is T / 2; If the predicted attenuation height is greater than zero and less than or equal to the first attenuation threshold, controlling the satellite to raise the predicted attenuation height between the first moment and the third moment, and starting the next adjustment cycle at the third moment; If the predicted attenuation height is greater than the first attenuation threshold and less than or equal to the second attenuation threshold, the satellite is controlled to raise the target height between the first moment and the third moment, and the next adjustment cycle is started at the third moment; The target height satisfies the following conditions: greater than or equal to the first attenuation threshold, less than or equal to the second attenuation threshold, and less than the first height, the first height being the elevation height of the satellite after continuous propulsion within T / 4 obtained based on the circular orbit formula, and the second attenuation threshold being twice the first attenuation threshold; Before controlling the satellite to raise the target altitude between the first moment and the third moment, the method further includes: Get the solar activity phase at the first moment; In the case where the solar activity phase is a quiet period, determining the first attenuation threshold as a target height; When the solar activity phase is at its peak, the minimum value between the second attenuation threshold and the first altitude is taken as the maximum elevation value, and the maximum elevation value is determined as the target altitude.

2. The method according to claim 1, characterized in that Also includes: When the solar activity phase is in a declining phase, the sum of the first attenuation threshold and the lifting adjustment amount is determined as the target height; When the solar activity phase is in the rising phase, the sum of the first attenuation threshold and twice the lifting adjustment amount is determined as the target height; The lift adjustment amount is the difference between the first attenuation threshold and the maximum lift value divided by three.

3. The method according to claim 1 or 2, characterized in that: After obtaining the predicted attenuation height of the satellite within the adjustment period, the method further includes: If the predicted attenuation height is greater than the second attenuation threshold and less than or equal to the third attenuation threshold, the satellite is controlled to raise the target height between the first moment and the fourth moment, and the next adjustment cycle is started at the fourth moment, wherein the third attenuation threshold is the average attenuation speed of the satellite during the solar flare period within the adjustment cycle, the third attenuation threshold is not less than four times the first attenuation threshold, and the interval between the first moment and the fourth moment is T / 4.

4. The method according to claim 3, characterized in that Before obtaining the predicted attenuation height of the satellite within the adjustment period, the method further includes: A plurality of average attenuation speeds of the satellite in a plurality of adjustment cycles before a first moment when the solar activity phase is at a peak period are obtained, and a maximum value among the plurality of average attenuation speeds is determined as the third attenuation threshold.

5. The method according to claim 1 or 2, characterized in that: After obtaining the predicted attenuation height of the satellite within the adjustment period, the method further includes: If the predicted attenuation height is greater than a third attenuation threshold, determining a first value n based on the predicted attenuation height and the target height; The satellite is controlled to raise the target altitude between the first moment and the fifth moment, and the next adjustment cycle is started at the fifth moment, wherein the interval between the first moment and the fifth moment is T / n.

6. The method according to claim 1, characterized in that Also includes: The target duration required for the satellite to rise to a specified altitude is calculated using the circular orbit formula, and the target duration is divided into multiple time periods to control the engine thrust to lift the satellite during the multiple time periods.

7. The method according to claim 1, characterized in that Before obtaining the predicted attenuation height of the satellite within the adjustment period, the method further includes: An average attenuation speed of the satellite in a previous adjustment period before the first moment is obtained, and the average attenuation speed in the previous adjustment period is determined as the first attenuation threshold.

8. An electronic device, characterized in that: The electronic device comprises: a processor and a memory storing computer program instructions; when the electronic device executes the computer program instructions, the method according to any one of claims 1 to 7 is implemented.

9. A computer program product, characterized in that The computer program product comprises computer program instructions, which implement the method according to any one of claims 1 to 7 when executed.

Images