Active orbital transfer method, device and product of ultra-low orbit satellite

Abstract

The invention discloses an active orbital transfer method, device and product of an ultra-low orbit satellite, and relates to the technical field of satellite detection. Comprising the steps of inputting a first instantaneous orbit element and an attitude strategy to an orbit mechanical model under the condition that the first instantaneous orbit element of a satellite at a first moment and the attitude strategy of the satellite in an adjustment period are obtained, and obtaining a second instantaneous orbit element of the satellite at a second moment; converting the second instantaneous orbital element into an average orbital element, and making a difference between a semi-major axis parameter in the average orbital element and a target average orbital semi-major axis to obtain a predicted attenuation height of the satellite in the adjustment period; if the predicted attenuation height is smaller than zero, the next adjustment period is started at the third moment, and the interval duration between the first moment and the third moment is T / 2; if the predicted attenuation height is larger than zero and smaller than or equal to the first attenuation threshold value, the satellite is controlled to rise the predicted attenuation height between the first moment and the third moment, and the next adjustment period is started 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 change method for an ultra-low orbit satellite, the method comprising: upon obtaining a first instantaneous orbital element of the satellite at a first moment and an attitude strategy of the satellite within an adjustment period, inputting the first instantaneous orbital element and the attitude strategy into an orbital mechanics model to obtain a second instantaneous orbital element of the satellite at a second moment, wherein 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 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, starting the next adjustment period at a 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 period at the third moment.

[0006] In a second aspect, an embodiment of the present application provides an active orbit change device for an ultra-low orbit satellite, the device comprising: an input module, for obtaining a first instantaneous orbital element of the satellite at a first moment, and an attitude strategy of the satellite in an adjustment period, inputting the first instantaneous orbital element and the attitude strategy into an orbital mechanics model to obtain a second instantaneous orbital element of the satellite at a second moment, wherein 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 length of the adjustment period is T; a determination module, for converting the second instantaneous orbital element into an average orbital element, and converting the average orbital element into an average orbital element. The semi-major axis parameter in the root number is subtracted from the target average orbit semi-major axis to obtain the predicted attenuation height of the satellite during the adjustment period, where the target average orbit semi-major axis is the average orbit semi-major axis corresponding to the standard altitude of the satellite; the orbit lifting module is used to start the next adjustment period at the third moment if the predicted attenuation height is less than zero, wherein the interval between the first moment and the third moment is T / 2; the orbit lifting module is also used to control the satellite to lift the predicted attenuation height between the first moment and the third moment if the predicted attenuation height is greater than zero and less than or equal to the first attenuation 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, comprising: a processor and a memory storing computer program instructions; when the processor executes the computer program instructions, the steps of the active orbit change method of the ultra-low orbit satellite as in the first aspect are implemented.

[0008] In a fourth aspect, 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 for an ultra-low orbit satellite as in the first aspect are implemented.

[0009] In a fifth aspect, 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, it implements the steps of the active orbit change method of the ultra-low orbit satellite as in the first aspect.

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

[0011] The present application provides an active orbit change method, device, equipment, medium and product for ultra-low orbit satellites. The semi-major axis of the average orbital element is used as a judgment index for satellite orbit attenuation. The attenuation height of the satellite in a future adjustment cycle can be predicted in advance to obtain the predicted attenuation height. When the satellite attenuation is predicted, the orbit raising operation is actively performed in advance to reduce the risk of falling. In this way, by predicting the orbit attenuation 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 application of the embodiments of the present application, the drawings in the embodiments of the present application are briefly introduced below.

[0013] Figure 1 It is a flowchart of an active orbit change method for an ultra-low orbit satellite provided in one embodiment of the present application; Figure 2 It is a structural schematic diagram of an active orbit change device for an ultra-low orbit satellite provided in an embodiment of the present application; Figure 3 It is a schematic diagram of the hardware structure of an electronic device provided in an embodiment of the present application. DETAILED DESCRIPTION

[0014] 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 implement the principles and spirit of the present application. The exemplary embodiments provided herein are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments herein, all other embodiments obtained by ordinary technicians of the art without creative work are within the scope of protection of this application.

[0015] In this document, terms such as first, second, third, etc. are only used to distinguish one entity (or operation) from another entity (or operation), but not to require or imply any order or relationship between these entities (or operations).

[0016] The following is a brief description of the concepts and technical terms that may be involved in the embodiments of the present application.

[0017] Instantaneous orbital elements: are orbital parameters that strictly satisfy the orbital equation, describing the instantaneous orbital state of the satellite at a specific moment. They include the influence of all perturbations (such as the non-spherical gravity of the earth, aerodynamic forces, gravity of the third celestial body, etc.). They describe the actual orbit of the satellite at this moment, including the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, perigee angle, and mean anomaly. Due to the existence of perturbations, the instantaneous orbital elements show periodic changes in time.

[0018] Mean orbital elements: orbital parameters obtained by eliminating the influence of short-period perturbations, usually used to describe the long-time orbital characteristics of satellites. Mean orbital elements can be derived from instantaneous orbital elements by filtering methods. Short-period perturbations are removed, and only long-term and medium-term perturbations are reflected. Changes are usually slow and more suitable for long-term analysis. It describes a smoothed orbit rather than the actual orbit of the satellite at a certain moment.

[0019] In related technologies, most satellite orbit management adopts a "post-response" strategy, that is, after the satellite orbit decays to a certain extent, the orbit raising operation is performed. The specific steps are as follows: ① Orbit monitoring: Real-time monitoring of the satellite’s orbital altitude and attenuation through ground measurement and control systems and onboard sensors.

[0020] ② Orbital decay warning: When it is monitored that the satellite orbit altitude drops below the preset safety threshold, an orbital decay warning will be issued.

[0021] ③ Orbit raising operation: Start the propulsion system on the satellite to perform orbit raising and restore the satellite to the predetermined orbital altitude.

[0022] However, the above orbit raising strategy has the problem of delayed response, that is, the orbit raising operation is carried out only after the orbit decays to a certain extent. 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 low position (high resistance, low thrust of ion thrusters), or the on-board power is insufficient (unable to push for a long time), the satellite will face a greater risk of crashing.

[0023] Based on this, this application can effectively avoid the risk of satellite crashes, improve the efficiency and reliability of orbit management, and solve the above-mentioned technical problems by predicting the orbital attenuation trend and proactively raising the orbit in advance.

[0024] The active orbit change method for an ultra-low orbit satellite provided in the embodiment of the present application is described in detail below with reference to the accompanying drawings through specific embodiments and their application scenarios.

[0025] Figure 1 It is a flow chart of an active orbit change method for an ultra-low orbit satellite provided in one embodiment of the present application. The executor of the active orbit change method for an ultra-low orbit satellite may be an orbit control system in an ultra-low orbit support system.

[0026] The following takes the orbit control system as an example to illustrate the active orbit change method of the ultra-low orbit satellite of the present application. It should be noted that the above-mentioned execution subjects and application scenarios do not constitute a limitation on the present application.

[0027] like Figure 1 As shown, the active orbit change method of the ultra-low orbit satellite provided in the embodiment of the present application may include steps 110 to 140.

[0028] Step 110, when a first instantaneous orbital element of the satellite at a first moment and an attitude strategy of the satellite within 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; Step 120, 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 the predicted attenuation height of the satellite within the adjustment period; Step 130, if the predicted attenuation height is less than zero, then starting the next adjustment cycle at the third moment; Step 140: If the predicted attenuation height is greater than zero and less than or equal to the first attenuation threshold, the satellite is controlled to raise the predicted attenuation height between the first moment and the third moment, and the next adjustment cycle is started at the third moment.

[0029] The active orbit change method for ultra-low orbit satellites provided in the embodiment of the present application uses the semi-major axis of the average orbital element as the judgment index of satellite orbit attenuation, and can predict the attenuation height of the satellite in a future adjustment cycle in advance, obtain the predicted attenuation height, and actively perform orbit raising operations in advance when the satellite attenuation is predicted to reduce the risk of falling. In this way, by predicting the orbit attenuation 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.

[0030] The specific implementation of the above steps will be described in detail below in conjunction with specific embodiments.

[0031] Involving step 110, when the first instantaneous orbital element of the satellite at the first moment and the attitude strategy of the satellite within the adjustment period are obtained, the first instantaneous orbital element and the attitude strategy are input into the orbital mechanics model to obtain the second instantaneous orbital element of the satellite at the second moment.

[0032] In step 110, the first instantaneous orbital element number is the satellite orbital element number of the satellite at the first moment, and the attitude strategy of the satellite during the adjustment period is the planned attitude parameters pre-input by relevant personnel. The attitude strategy is used to control the operating attitude of the satellite during the adjustment period, including the pitch angle, roll angle, and yaw angle.

[0033] 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 element is the satellite orbital element at the second moment, which is a predicted value calculated by the orbital mechanics model combined with the first instantaneous orbital element and attitude strategy.

[0034] Satellite orbit elements are a set of six key parameters needed to describe the motion of a celestial body or spacecraft in its Kepler orbit, including 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.

[0035] The orbital mechanics model may be a simplified general perturbation model 4 (SGP4), a high precision orbit estimator / high precision orbit prediction program (HPOP), etc., and this application does not make any specific limitation on this.

[0036] It should be noted that the prediction cycle of the orbital mechanics model is shorter than the adjustment cycle. The orbital mechanics model does not directly predict the second instantaneous orbital element based on the first instantaneous orbital element, but needs to make multiple predictions to get it. For example, if the prediction cycle is 3 hours and the adjustment cycle is 24 hours, the first prediction will get the instantaneous orbital element of the 3rd hour, and the second prediction will get the instantaneous orbital element of the 6th hour (predicted value). After multiple predictions, the instantaneous orbital element of the 24th hour is obtained based on the instantaneous orbital element of the 21st hour (predicted value), which is the second instantaneous orbital element.

[0037] The adjustment period has a duration of T, and is a period used to predict the attenuation height of the satellite and adjust the real-time operating height of the satellite based on the attenuation height.

[0038] Because the lower the satellite's altitude, the greater the risk of crashing under the passive orbit raising strategy, the duration of this cycle is positively correlated with the satellite's working altitude (or operating altitude). That is, the lower the satellite's altitude, the shorter the adjustment cycle. In other words, the satellite's real-time operating altitude is adjusted with a higher adjustment frequency. When satellite attenuation is predicted, the orbit raising operation can be performed in time in advance to reduce the risk of crashing.

[0039] For example, if the satellite is operating at an altitude of 300 km, the recommended value of the adjustment period T is 24 hours. As the satellite operating altitude decreases, the T value needs to be reduced.

[0040] Involving step 120, the second instantaneous orbital element is converted into an average orbital element, and the semi-major axis parameter in the average orbital element is subtracted from the target average orbital semi-major axis to obtain the predicted attenuation height of the satellite within the adjustment period.

[0041] In step 120, the semi-major axis parameter in the second instantaneous orbital element is the instantaneous orbital semi-major axis, and the semi-major axis parameter in the average orbital element is the instantaneous average orbital semi-major axis. The present application can use a Fourier transform algorithm or a time smoothing algorithm to perform time averaging on the instantaneous orbital semi-major axis in the second instantaneous orbital element to obtain the instantaneous average orbital semi-major axis corresponding to the satellite at the second moment. The predicted attenuation height is the predicted value of the attenuation height of the satellite after one adjustment cycle.

[0042] For example, at the start time t0 of the adjustment cycle, the first instantaneous orbital element of the satellite is obtained through telemetry data, and the attitude strategy of the satellite in this cycle is used as input to perform orbital extrapolation of a cycle length T on the satellite to obtain the second instantaneous orbital element of the satellite at time t0+T, and the result is converted to obtain the instantaneous average orbital semi-major axis , then the predicted attenuation height , is the target mean orbit semi-major axis.

[0043] Converting instantaneous orbital elements to the semi-major axis of the average orbital elements is a complex process, because instantaneous orbital elements (such as position, velocity, etc.) vary with time, while average orbital elements (such as average semi-major axis, average eccentricity, etc.) are obtained by averaging over time. This conversion usually involves knowledge of orbital mechanics and perturbation theory. In orbital mechanics, instantaneous orbital elements usually include six parameters: semi-major axis, eccentricity, inclination, right ascension of ascending node, argument of perigee and true anomaly. These parameters describe the orbital state of the satellite at a certain moment. However, due to factors such as the non-sphericity of the earth, atmospheric drag, and gravitational perturbations of the sun and the moon, the orbit of the satellite will change over time. Therefore, the instantaneous orbital elements will also change over time. In order to obtain the average orbital elements, the instantaneous orbital elements need to be averaged over time. This usually involves numerically integrating the satellite orbit and calculating the average value of each orbital element within a certain time interval.

[0044] It should be noted that this application only uses the semi-major axis parameter in the average orbital element when calculating the predicted attenuation height. Therefore, it is only necessary to convert the semi-major axis parameter in the second instantaneous orbital element into the instantaneous average orbital semi-major axis, and there is no need to convert the remaining five parameters.

[0045] In some embodiments of the present application, the period duration of the oscillation cycle is T0, and the instantaneous orbital semi-major axis in the second instantaneous orbital element is time-averaged by Fourier transform to obtain the instantaneous average orbital semi-major axis corresponding to the satellite at the second moment, which can specifically include: obtaining all instantaneous orbital elements of the satellite within at least one oscillation cycle before the second moment to obtain a data set; performing Fourier transform on the second instantaneous orbital element and all instantaneous orbital semi-major axes in the data set to obtain a frequency domain representation of the orbital semi-major axis; filtering the frequency domain representation based on a low-pass filter to truncate high-frequency components, retain only DC components and low-frequency components (representing average values ​​and long-term changes), and filtering out high-frequency periodic oscillations; performing an inverse Fourier transform on the filtered frequency domain representation to obtain a filtered time series; and using the filtered time series as the instantaneous average orbital semi-major axis of the satellite at the second moment.

[0046] In some other embodiments of the present application, the period duration of the oscillation period is T0, and the instantaneous orbital semi-major axis in the second instantaneous orbital element is time-averaged using a time smoothing algorithm to obtain the instantaneous average orbital semi-major axis corresponding to the satellite at the second moment. This may specifically include: taking the second moment as the intermediate moment and the period duration T0 as the time window length to determine the target time window; obtaining all instantaneous orbital elements of the satellite in the target time window, and performing a sliding average calculation on the average value of the instantaneous orbital semi-major axis in all instantaneous orbital elements in the target time window to obtain the instantaneous average orbital semi-major axis of the satellite at the second moment.

[0047] It should be noted that when all instantaneous orbital elements of the satellite in the target time window are obtained, the orbital mechanics model also needs to predict the instantaneous orbital elements in the time period T0 / 2 after the second moment.

[0048] The mean orbital semi-major axis of a satellite is an important parameter that determines the shape and period of the satellite's orbit around the Earth. is the average orbital semi-major axis corresponding to the standard altitude h of the satellite (i.e. the average altitude of the satellite from the earth's surface). The standard altitude of the satellite 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.

[0049] For example, h is the average altitude of 300 km, then =6371+300=6671km.

[0050] In step 130 , if the predicted attenuation height is less than zero, the next adjustment cycle is started at the third moment.

[0051] In step 130, the interval between the first moment and the third moment is T / 2. If the predicted attenuation height is less than zero, it indicates that the satellite is temporarily at no risk of falling in the next adjustment cycle. Therefore, the orbit raising operation is not performed on the satellite, and the next adjustment cycle is started T / 2 in advance. The future attenuation height of the satellite is predicted at the beginning of the next adjustment cycle, and it is determined whether to perform the orbit raising operation based on the future attenuation height.

[0052] Exemplarily, T is 24 hours, and the first moment is 0:01 on February 1st. If the predicted attenuation altitude of the satellite from 0:01 on February 1st to 24:00 on February 1st is less than zero, the next adjustment cycle can be started at 12:01 on February 1st, and the attenuation altitude of the satellite in the next 24 hours can be predicted at 12:01 on February 1st.

[0053] Involving step 140, if the predicted attenuation height is greater than zero and less than or equal to the first attenuation threshold, the satellite is controlled to raise the predicted attenuation height between the first moment and the third moment, and the next adjustment cycle is started at the third moment.

[0054] In step 140, the first attenuation threshold may be pre-calculated or pre-set, and this application does not make any specific limitation on this.

[0055] It should be noted that the above steps 130 and 140 are two judgment branches, so this application does not specifically limit their execution order.

[0056] In some embodiments of the present application, the first attenuation threshold may be obtained by precalculation. Before obtaining the predicted attenuation height of the satellite within the adjustment period in step 120, the following steps may also be included: 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 a first attenuation threshold.

[0057] In the embodiment of the present application, the satellite orbit decay speed is affected by the solar activity cycle, which is about 11 years and is divided into high solar activity years and low solar activity years. In high solar activity years, solar eruptions are frequent, the atmospheric density increases, and the satellite orbit decay speed is significantly accelerated. Therefore, the present application can determine the average decay speed of the satellite in the last adjustment cycle before the first moment as the first decay threshold. In this way, the first decay threshold It can change with the changes in solar activity. For example, for a satellite at an altitude of 300km, the attenuation rate is about 1-2km / day during the solar calm period, and the attenuation rate can reach 5km / day during the solar flare period, so that the size of the first attenuation threshold can be more in line with the operating status of the satellite under the current solar activity state.

[0058] In some embodiments of the present application, in order to accurately lift the satellite to a specified orbital altitude, when performing an orbital lifting operation on the satellite, it can include: calculating the target time required to lift the satellite to the specified altitude through a circular orbit formula, and dividing the target time into multiple time periods, so as to control the engine to start thrust to lift the satellite within multiple time periods.

[0059] Specifically, according to the circular orbit formula, the tangential velocity of the satellite in orbit is , then we want to change the semi-major axis to The satellite in the circular orbit is raised to a semi-major axis of The change in linear velocity on a circular track , the target duration of the launch can be further calculated based on the satellite's engine thrust F and satellite mass m: . After that, it can be divided into multiple time periods, and the sum of the duration is t.

[0060] In an embodiment of the present application, the circular orbit formula can be used to accurately calculate the target duration required for the satellite to rise to a specified altitude. Based on this, the target duration is divided into multiple time periods, and the engine is controlled to start thrust during multiple time periods, so that the satellite can be accurately lifted to the specified orbital altitude.

[0061] In other embodiments of the present application, in order to accurately lift the satellite to a specified altitude, when performing an orbital lifting operation on the satellite, it can include: while the engine is continuously thrusting, converting the instantaneous orbital elements fed back by the satellite in real time into average orbital elements, and determining the actual lift height of the satellite based on the semi-major axis parameters in the average orbital elements, and continuing to thrust when the actual lift height is less than the specified altitude until the actual lift height reaches the specified altitude.

[0062] In an embodiment of the present application, the engine can be controlled to continuously thrust, and during the thrust period the actual lift 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 that the thrust is continued when the actual lift height has not reached the specified height, thereby achieving a one-time lift of the satellite to the specified height without stopping in between, and the orbit lift speed is faster and more efficient.

[0063] In some embodiments of the present application, in order to timely remedy the satellite when the satellite attenuation degree is high, after obtaining the predicted attenuation height of the satellite within the adjustment period in the above step 120, the following steps may also be included: Step 150: 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.

[0064] 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 , less than the first altitude, the first altitude is the altitude of the satellite after continuous propulsion within T / 4 obtained based on the circular orbit formula , the second attenuation threshold is twice the first attenuation threshold.

[0065] In the embodiment of the present application, the first attenuation threshold is the average attenuation speed of the satellite in the last adjustment cycle before the first moment, so the first attenuation threshold is set to the minimum value of the target height, so that the satellite's lifting height matches the satellite's attenuation speed in the adjustment cycle. If the actual attenuation height does not reach the predicted value, the satellite will be lifted too high and will deviate from its original orbit (at an orbital altitude of 300km, a height difference of 2km will cause the satellite position to deviate by an order of magnitude of 100 kilometers per day). Therefore, by limiting the target altitude to be less than or equal to twice the first attenuation threshold, the satellite's lifting height is restricted to prevent the satellite from being lifted too high. Combined with the satellite's lifting capacity design, it is less than the height that the satellite is lifted by continuously pushing during a quarter of the adjustment cycle, so that the satellite's lifting height meets the engine capacity, and the target altitude is not set too high and exceeds the engine's operating capacity.

[0066] In this way, a strategy of small-amplitude lifting and high-frequency prediction is adopted. High frequency means that the cycle of the lifting operation is shorter than the entire adjustment cycle T, and the frequency decreases as the attenuation risk increases (T / 2, T / 4 or less). The advantage of this is that when the orbital attenuation forecast is too small or too large, the next adjustment cycle can be started in time, thereby reducing the risk caused by the forecast error; small-amplitude lifting means that the height of each lift is no higher than twice the first attenuation threshold, which avoids the orbital deviation caused by the satellite being lifted too high when the predicted attenuation is too large relative to the actual attenuation. When the actual attenuation risk is too large, the high frequency makes up for the lack of lifting amplitude. In summary, this target height setting strategy increases the robustness of the overall orbit lifting scheme and reduces the risk of orbit prediction errors.

[0067] In related technologies, when encountering extreme space weather such as solar flares and geomagnetic explosions, the atmospheric density will change significantly, causing a sudden increase in the air resistance of ultra-low orbit satellites. If the satellite is at a relatively low position at this time, it will not be able to propel for a long time due to the large air resistance and small thrust of the ion thruster, which will lead to a greater risk of the satellite crashing.

[0068] In some embodiments of the present application, before the above-mentioned step 150 controls the satellite to raise the target altitude between the first moment and the third moment, it may also include: obtaining the solar activity stage at the first moment; when the solar activity stage is a calm period, determining the first attenuation threshold as the target altitude; when the solar activity stage is a peak period, taking the minimum value of the second attenuation threshold and the first altitude as the maximum elevation value, and determining the maximum elevation value as the target altitude; when the solar activity stage is a declining period, determining the sum of the first attenuation threshold and the elevation adjustment amount as the target altitude; when the solar activity stage is an rising period, determining the sum of the first attenuation threshold and twice the elevation adjustment amount as the target altitude.

[0069] The lift adjustment amount is the difference between the first attenuation threshold and the lift maximum value divided by three.

[0070] In the embodiment of the present application, the height of the target altitude is related to the intensity of solar activity. The more intense the solar activity, the higher the target altitude. In this way, when the solar activity is relatively calm, the air resistance of the ultra-low orbit satellite is relatively small, the risk of the satellite crashing is low, and the safety factor is high. At this time, a relatively low lift altitude can be set for the satellite within a reasonable range to avoid the satellite from leaving the original orbit due to excessive lift; when the solar activity is more intense, the air resistance of the ultra-low orbit satellite is relatively large, the risk of the satellite crashing is high, and the safety factor is low. At this time, a relatively high lift altitude can be set for the satellite within a reasonable range to avoid the situation where the satellite is in a low position and cannot be pushed for a long time due to large air resistance and small thrust of the ion thruster, thereby reducing the risk of the satellite crashing.

[0071] In some embodiments of the present application, after obtaining the predicted attenuation height of the satellite within the adjustment period in the above step 120, it may also include: if the predicted attenuation height is greater than the second attenuation threshold and less than or equal to the third attenuation threshold, controlling the satellite to raise the target height between the first moment and the fourth moment, and starting the next adjustment period at the fourth moment, wherein the third attenuation threshold is the average attenuation speed of the satellite during the solar flare period during the adjustment period, 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.

[0072] In an embodiment of the present application, the third attenuation threshold is the average attenuation speed of the satellite during the solar flare period within the adjustment period. Therefore, the third attenuation threshold represents a higher degree of attenuation. If the predicted attenuation height of the satellite is between the second attenuation threshold and the third attenuation threshold, it indicates that the attenuation degree of the satellite during the adjustment period is more serious. Based on this, the present application further shortens the orbit raising time of the satellite, speeds up the orbit raising process of the satellite, and quickly raises the satellite to the target height within T / 4 after the first moment. In addition, the next adjustment period is started after T / 4, so that the period of the lifting operation is shorter than the entire adjustment period T. When the orbit attenuation forecast is too large, the next adjustment period can be started in time, thereby reducing the risk caused by the forecast error.

[0073] In other embodiments of the present application, before obtaining the predicted attenuation height of the satellite within the adjustment period in the above-mentioned step 120, it may also include: obtaining multiple average attenuation speeds of the satellite in multiple adjustment periods before the first moment when the solar activity phase is at its peak, and determining the maximum value of the multiple average attenuation speeds as the third attenuation threshold.

[0074] In some embodiments of the present application, after obtaining the predicted attenuation height of the satellite within the adjustment period in the above step 120, it may also include: if the predicted attenuation height is greater than the third attenuation threshold, determining a first value n based on the predicted attenuation 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.

[0075] Specifically, , is the target height, and the interval between the first moment and the fifth moment is T / n.

[0076] In the embodiment of the present application, if the predicted attenuation height of the satellite is greater than the third attenuation threshold, it indicates that the attenuation degree of the satellite in the adjustment period is serious, so it is necessary to further shorten the time spent on the satellite's orbital elevation. By dividing the predicted attenuation height by the target height and rounding it up and adding 1, the satellite is quickly raised to the target height within T / n after the first moment, further accelerating the satellite's orbital elevation process. In addition, the next adjustment period is started after T / n, so that the period of the lifting operation is shorter than the entire adjustment period T. When the orbital attenuation forecast is too large, the next adjustment period can be started in time, thereby reducing the risk caused by the forecast error.

[0077] As a specific example, T = 24h, the predicted attenuation height of the satellite during the adjustment period is , the first attenuation threshold, the second attenuation threshold, and the third attenuation threshold are , the target height is , then the track lifting strategy provided in this application is as follows: ① , indicating that the algorithm predicts that the satellite's altitude after one cycle will be above the standard altitude. At this time, the satellite does not need to perform an elevation operation. Always start the next adjustment cycle; ② , indicating that the predicted attenuation altitude is slightly lower than the standard altitude. At this time, the satellite needs to be lifted within 12 hours. , and in Always start the next adjustment cycle; ③ , the satellite needs to be lifted within 12 hours , and in Always start the next adjustment cycle; ④ , the satellite needs to be lifted within 6 hours , and in Always start the next adjustment cycle; ⑤ , indicating that the predicted attenuation height will significantly attenuate the orbit in one cycle. Human intervention can be added at this time. If there is no human intervention, the adjustment frequency needs to be further increased. , , the satellite needs to be Lifting within time , and in Always start the next adjustment cycle.

[0078] In some embodiments of the present application, the accurate calculation of the predicted attenuation 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 orbital correction model. The specific training steps of the above-mentioned orbital correction model are as follows: Acquire multiple time features, satellite state features, space environment features and satellite orbit elements at the first moment; input the satellite orbit elements at the first moment into the orbital mechanics model to obtain the predicted value of the velocity and position parameters at the second moment output by the orbital mechanics model, wherein the first moment and the second moment are separated by a preset period; construct historical training data by combining the time features, satellite state features, space environment features of each first moment, and the predicted value and true value of the velocity and position parameters at the second moment; use the time features, satellite state features, space environment features of the first moment, and the predicted value of the velocity and position parameters at the second moment as input training data, and use the difference between the predicted value and the true value of the velocity and position parameters at the second moment as output training data (i.e., output label) for training.

[0079] 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 value of the speed and position parameters at the second moment can be obtained by actually detecting the satellite through the satellite detection system. The mean square error can be used as the loss function during model training, and the back propagation algorithm is used for training. The orbit correction model is associated with the orbital mechanics model. The preset model can be set according to specific needs, and this application does not make specific restrictions on this. The difference between the predicted value and the true value of the speed and position parameters at the second moment can be the true error value. The mean square error of the training error prediction value output by the preset model and the true error value is calculated to obtain the loss function value.

[0080] It should be noted that in the historical training data, the satellite state characteristics are in the ECEF coordinate system, the space environment characteristics are in the geodetic coordinate system, and the predicted value and true value 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 earth-fixed coordinate system.

[0081] In an embodiment of the present application, during the training process of the orbit correction model, the training data used is derived from the predicted values ​​actually output by the orbital mechanics model. Based on this, during the training process, the orbit correction model can learn the deviation between the predicted values ​​output by the orbital mechanics model under different time characteristics, different satellite state characteristics, and different space environment characteristics and the true values, so that the trained orbit correction model can have the ability to reasonably and accurately evaluate the error of the predicted value output by the orbital mechanics model in different satellite orbit prediction scenarios by combining time characteristics, satellite state characteristics, and space environment characteristics, and obtain an orbit correction model that is compatible with the orbital mechanics model, so that the orbit correction model can have better prediction error evaluation capabilities for the associated orbital mechanics model.

[0082] Corresponding to the method embodiment of the present application, the present application also provides an active orbit change device for an ultra-low orbit satellite.

[0083] Figure 2 Schematic diagram of the structure of an active orbit change device for an ultra-low orbit satellite provided in an embodiment of the present application. Figure 2 As shown, the active orbit change device 200 of the ultra-low orbit satellite may include: an input module 210, a determination module 220, and an orbit raising module 230.

[0084] Among them, the input module 210 is used to input the first instantaneous orbital element and the attitude strategy into the orbital mechanics model after obtaining the first instantaneous orbital element of the satellite at the first moment and the attitude strategy of the satellite in the adjustment period, so as to obtain the second instantaneous orbital element of the satellite at the second moment, wherein the first moment is the starting moment of the adjustment period, the second moment is the ending moment of the adjustment period, and the period length of the adjustment period is T; the determination module 220 is used to convert the second instantaneous orbital element into the average orbital element, and to subtract the semi-major axis parameter in the average orbital element from the target average orbital semi-major axis to obtain the predicted attenuation height of the satellite in the adjustment period, and the target average orbital semi-major axis is the average orbital semi-major axis corresponding to the standard altitude of the satellite operation; the orbit lifting module is used to start the next adjustment period at the third moment if the predicted attenuation height is less than zero, wherein the interval between the first moment and the third moment is T / 2; the orbit lifting module 230 is also used to control the satellite to lift the predicted attenuation height between the first moment and the third moment if the predicted attenuation height is greater than zero and less than or equal to the first attenuation threshold, and to start the next adjustment period at the third moment.

[0085] The active orbit change device for ultra-low orbit satellites provided in the embodiment of the present application uses the semi-major axis of the average orbital element as a judgment index for satellite orbit attenuation, and can predict in advance the attenuation height of the satellite in a future adjustment cycle to obtain the predicted attenuation height. When the satellite attenuation is predicted, the orbit raising operation is actively performed in advance to reduce the risk of falling. In this way, by predicting the orbit attenuation 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.

[0086] In some embodiments of the present application, the orbit raising module is also used to, after obtaining the predicted attenuation height of the satellite within the adjustment period, control the satellite to raise the target height between the first moment and 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, 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 attenuation threshold, less than or equal to the second attenuation threshold, less than the first height, the first height being the elevation height of the satellite after continuous push within T / 4 obtained based on the circular orbit formula, and the second attenuation threshold being twice the first attenuation threshold.

[0087] In some embodiments of the present application, the orbital lifting module is also used to: obtain the solar activity stage at the first moment; when the solar activity stage is a calm period, determine the first attenuation threshold as the target altitude; when the solar activity stage is a peak period, take the minimum value of the second attenuation threshold and the first altitude as the maximum lifting value, and determine the maximum lifting value as the target altitude; when the solar activity stage is a declining period, determine the sum of the first attenuation threshold and the lifting adjustment amount as the target altitude; when the solar activity stage is an rising period, determine the sum of the first attenuation threshold and twice the lifting adjustment amount as the target altitude; wherein the lifting adjustment amount is the difference between the first attenuation threshold and the maximum lifting value divided by three.

[0088] In some embodiments of the present application, the orbit raising module is also used for: after obtaining the predicted attenuation height of the satellite within the adjustment period, if the predicted attenuation height is greater than the second attenuation threshold and less than or equal to the third attenuation threshold, then controlling the satellite to raise the target height between the first moment and the fourth moment, and starting the next adjustment period at the fourth moment, wherein the third attenuation threshold is the average attenuation speed of the satellite during the solar flare period during the adjustment period, 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.

[0089] In some embodiments of the present application, the determination module is also used to obtain multiple average attenuation speeds of the satellite in multiple adjustment periods before a first moment when the solar activity phase is at a peak before obtaining the predicted attenuation height of the satellite in the adjustment period, and determine the maximum value of the multiple average attenuation speeds as the third attenuation threshold.

[0090] In some embodiments of the present application, the orbit raising module is also used for: after obtaining the predicted attenuation height of the satellite within the adjustment period, if the predicted attenuation height is greater than the third attenuation threshold, determining a first value n based on the predicted attenuation 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, wherein the interval between the first moment and the fifth moment is T / n.

[0091] In some embodiments of the present application, the orbit raising module is also used to: calculate the target time required to raise the satellite to a specified height through a circular orbit formula, and divide the target time into multiple time periods to control the engine to start thrust to raise the satellite within multiple time periods.

[0092] In some embodiments of the present application, the determination module is also used to obtain the average attenuation speed of the satellite in the previous adjustment period before the first moment before obtaining the predicted attenuation height of the satellite in the adjustment period, and determine the average attenuation speed in the previous adjustment period as the first attenuation threshold.

[0093] The active orbit change device for ultra-low orbit satellites provided in the embodiment of the present application can achieve Figure 1 The various processes implemented by the service platform in the method embodiment can achieve the same technical effect. To avoid repetition, they will not be described here.

[0094] Figure 3 It is a schematic diagram of the hardware structure of an electronic device provided in an embodiment of the present application.

[0095] like Figure 3 As shown, the electronic device 300 includes a memory 301 , a processor 302 , and a computer program stored in the memory 301 and executable on the processor 302 .

[0096] In an example, the processor 302 may include a central processing unit (CPU), or an application specific integrated circuit (ASIC), or may be configured to implement one or more integrated circuits of the embodiments of the present application.

[0097] The memory 301 may include a read-only memory (ROM), a random access memory (RAM), a disk storage medium device, an optical storage medium device, a flash memory device, an electrical, optical or other physical / tangible memory storage device. Therefore, generally, the memory includes one or more tangible (non-transitory) computer-readable storage media (e.g., a memory device) 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 according to the first aspect of the present application.

[0098] The processor 302 runs a computer program corresponding to the executable program code by reading the executable program code stored in the memory 301, so as to implement the method in the embodiment of the first aspect.

[0099] In some examples, the electronic device 300 may further include a communication interface 303 and a bus 310. Figure 3 As shown, the memory 301, the processor 302, and the communication interface 303 are connected via a bus 310 and communicate with each other.

[0100] The communication interface 303 is mainly used to realize the communication between the modules, devices, units and / or equipment in the embodiment of the present application. The communication interface 303 can also be used to access input devices and / or output devices.

[0101] The bus 310 includes hardware, software, or both, coupling the components of the electronic device 300 to each other. By way of example and not limitation, the bus 310 may 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 may include one or more buses. Although embodiments of the present application describe and illustrate a particular bus, the present application contemplates any suitable bus or interconnect.

[0102] The electronic device provided in the embodiment of the present application can realize Figure 1 The various processes implemented by the electronic device in the method embodiment can achieve the same technical effect, and to avoid repetition, they will not be described here.

[0103] In combination with the active orbit change method of the ultra-low orbit satellite in the above embodiment, the embodiment of the present application can provide a computer storage medium for implementation. The computer storage medium stores computer program instructions; when the computer program instructions are executed by the processor, the steps of any one of the active orbit change methods of the ultra-low orbit satellite in the above embodiment are implemented.

[0104] In combination with the active orbit change method of the ultra-low orbit satellite in the above embodiment, the embodiment of the present application can provide a computer program product for implementation. 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 embodiment are implemented.

[0105] An embodiment of the present application further provides a chip, which includes a processor and a communication interface, wherein the communication interface and the processor are coupled, and the processor is used to run programs or instructions to implement the various processes of the above-mentioned active orbit change method embodiment of the ultra-low orbit satellite, and can achieve the same technical effect. To avoid repetition, it will not be repeated here.

[0106] It should be understood that the chip mentioned in the embodiments of the present application can also be called a system-level chip, a system chip, a chip system or a system-on-chip chip, etc.

[0107] It should be clear that the present application is not limited to the specific configuration and processing described above and shown in the figures. For the sake of simplicity, a detailed description of the known method 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, and those skilled in the art can make various changes, modifications and additions, or change the order between the steps after understanding the spirit of the present application.

[0108] The functional blocks shown in the structural block diagram described above 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 function card, etc. 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 on a transmission medium or a communication link via a data signal carried in a carrier. "Machine-readable medium" may include any medium capable of storing or transmitting information. Examples of machine-readable media include electronic circuits, semiconductor memory devices, ROMs, flash memories, erasable ROMs (EROMs), floppy disks, CD-ROMs, optical disks, hard disks, optical fiber media, radio frequency (RF) links, etc. The code segment can be downloaded via a computer network such as the Internet, an intranet, etc.

[0109] It should also be noted that the exemplary embodiments mentioned in this application describe some methods or systems based on a series of steps or devices. However, this application is not limited to the order of the above steps, that is, the steps can be performed in the order mentioned in the embodiment, or in a different order from the embodiment, or several steps can be performed simultaneously.

[0110] Aspects of the present disclosure are described above with reference to the flowcharts and / or block diagrams of the methods, devices (systems) and computer program products according to the embodiments of the present disclosure. It should be understood that each box in the flowchart and / or block diagram and the combination of each box 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 device to produce a machine so that these instructions executed by the processor of the computer or other programmable data processing device enable the implementation of the functions / actions specified in one or more boxes 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 can also be understood that each box in the block diagram and / or flowchart and the combination of boxes in the block diagram and / or flowchart can also be implemented by dedicated hardware that performs the specified function or action, or can be implemented by a combination of dedicated hardware and computer instructions.

[0111] The above is only a specific implementation of the present application. Those skilled in the art can clearly understand that for the convenience and simplicity of description, the specific working processes of the systems, modules and units described above can refer to the corresponding processes in the aforementioned method embodiments, and will not be repeated here. It should be understood that the protection scope of the present application is not limited to this. Any technician familiar with the technical field can easily think of various equivalent modifications or replacements within the technical scope disclosed in the present application, and these modifications or replacements should be included in 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 ; If the predicted attenuation height is greater than zero and less than or equal to the first attenuation threshold, the satellite is controlled to raise the predicted attenuation height between the first moment and the third moment, and the next adjustment cycle is started at the third moment.

2. The method according to claim 1, 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 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, where the first height is the height of the satellite obtained based on the circular orbit formula. The lifting height after continuous pushing, the second attenuation threshold is twice the first attenuation threshold.

3. The method according to claim 2, characterized in that 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 a peak, taking the minimum value between the second attenuation threshold and the first height as the maximum elevation value, and determining the maximum elevation value as the target height; 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.

4. The method according to any one of claims 2 to 3, 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 .

5. The method according to claim 4, 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.

6. The method according to any one of claims 2 to 3, 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 .

7. 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.

8. 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.

9. 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 8 is implemented.

10. 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 8 when executed.

Images