Ultra-low Orbit Determination Method, Device, Medium and Product for Low Earth Orbit Satellites

Abstract

The present application discloses a method, device, medium and product for determining the ultra-low orbit of a near-earth satellite. The method includes: when the thrust of the ion thruster of the satellite is constant, obtaining the thrust vector of the satellite engine in the satellite body coordinate system; obtaining the first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and converting the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain the thrust components of the thrust vector on the three coordinate axes of the geocentric inertial coordinate system; combining the satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes to obtain a small thrust perturbation term; adding the small thrust perturbation term on the basis of the atmospheric drag perturbation term and the conventional drag perturbation term to obtain an input parameter, and inputting the input parameter into the orbital mechanics model, so that the orbital mechanics model determines the satellite orbit elements based on the atmospheric drag perturbation term, the conventional perturbation term and the small thrust perturbation term.

Description

Technical Field

[0001] The present application relates to the technical field of space environment detection, particularly to the technical field of satellite detection, and specifically to a method, device, medium and product for determining the ultra-low orbit of a near-earth satellite. Background Art

[0002] Ultra-low orbit satellites operate on orbits between 150 km and 300 km from the Earth's surface. Due to factors such as a large number of satellites and an increased impact of the atmospheric environment, ultra-low orbit satellites now face challenges such as rapid orbital decay, increased collision risks, and increased difficulty in attitude control. Therefore, it is crucial to build a complete ultra-low orbit guarantee system to ensure the stable operation of satellites and the efficient execution of tasks.

[0003] Currently, when predicting the future orbital state of a satellite, existing orbital calculation methods mainly rely on the atmospheric drag perturbation term and the conventional perturbation term. This calculation method has obvious limitations when facing high-precision requirements, resulting in a certain deviation between the calculated satellite orbital elements and the actual orbital state. Especially in low orbits and high-precision tasks, it cannot meet the high-precision requirements for the satellite orbital position. Summary of the Invention

[0004] In view of this, embodiments of the present application provide a method, device, equipment, medium and product for determining the ultra-low orbit of a near-earth satellite, which can improve the prediction accuracy of the satellite orbital position.

[0005] In a first aspect, embodiments of the present application provide a method for determining the ultra-low orbit of a near-earth satellite. The method includes: when the thrust of the ion thruster of the satellite is constant, obtaining the thrust vector of the satellite engine in the satellite body coordinate system; obtaining the first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and based on the first rotation matrix, converting the thrust vector to the geocentric inertial coordinate system to obtain the thrust components of the thrust vector on the three coordinate axes of the geocentric inertial coordinate system; combining the satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes to obtain a small thrust perturbation term; adding the small thrust perturbation term to the atmospheric drag perturbation term and the conventional drag perturbation term to obtain input parameters, and inputting the input parameters into an orbital mechanics model, so that the orbital mechanics model determines the satellite orbital elements based on the atmospheric drag perturbation term, the conventional perturbation term, and the small thrust perturbation term.

[0006] Second aspect, an ultra-low orbit determination device for a near-earth satellite provided by an embodiment of the present application includes: an acquisition module, configured to acquire the thrust vector of the satellite engine in the satellite body coordinate system when the thrust magnitude of the ion thruster of the satellite is constant; a coordinate system conversion module, configured to acquire the first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and convert the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain the thrust components of the thrust vector on the three coordinate axes of the geocentric inertial coordinate system; a combination module, configured to combine the satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes to obtain a small thrust perturbation term; a determination module, configured to add the small thrust perturbation term on the basis of the atmospheric drag perturbation term and the conventional drag perturbation term to obtain an input parameter, and input the input parameter into an orbital mechanics model, so that the orbital mechanics model determines the satellite orbit elements based on the atmospheric drag perturbation term, the conventional perturbation term, and the small thrust perturbation term.

[0007] Third aspect, an electronic device provided by an embodiment of the present application includes: a processor and a memory storing computer program instructions; when the processor executes the computer program instructions, the steps of the ultra-low orbit determination method for a near-earth satellite as in the first aspect are implemented.

[0008] Fourth aspect, a computer-readable storage medium provided by an embodiment of the present application stores computer program instructions, and when the computer program instructions are executed by a processor, the steps of the ultra-low orbit determination method for a near-earth satellite as in the first aspect are implemented.

[0009] Fifth aspect, a computer program product provided by an embodiment of the present application is stored in a non-volatile storage medium, and when the computer program product is executed by a processor, the steps of the ultra-low orbit determination method for a near-earth satellite as in the first aspect are implemented.

[0010] Sixth aspect, a chip provided by an embodiment of the present application includes a processor and a communication interface, the communication interface is coupled to the processor, and the processor is configured to run a program or instructions to implement the steps of the ultra-low orbit determination method for a near-earth satellite as in the first aspect.

[0011] The present application provides a method, apparatus, device, medium and product for determining the ultra-low orbit of a near-earth satellite. Considering that in the previous satellite orbit position calculation scheme, only the atmospheric drag perturbation term and the conventional perturbation term were considered, and the influence of the small thrust perturbation term on the change of the satellite orbit position was not considered. Based on this, the present application adds a small thrust perturbation term to the input parameters of the orbit mechanics model, and combines the small thrust perturbation term, the atmospheric drag perturbation term and the conventional perturbation term to more comprehensively reflect the actual force condition of the satellite orbit, so that the orbit mechanics model can analyze the orbit change of the satellite under the action of various forces according to the actual force condition, and more accurately determine the satellite orbit elements, making the calculated orbit elements closer to the actual orbit state. Since the small thrust perturbation term is considered, the calculation of the satellite position and velocity during the integration process is more accurate, better meeting the high-precision requirements of modern space missions for orbit position calculation. 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 flow chart of a method for determining the ultra-low orbit of a near-earth satellite provided by an embodiment of the present application;

[0014] Figure 2 It is a schematic flow chart of a method for determining the ultra-low orbit of a near-earth satellite provided by another embodiment of the present application;

[0015] Figure 3 It is a schematic flow chart of a method for determining the ultra-low orbit of a near-earth satellite provided by still another embodiment of the present application;

[0016] Figure 4 It is a schematic structural diagram of an apparatus for determining the ultra-low orbit of a near-earth satellite provided by an embodiment of the present application;

[0017] Figure 5 It is a schematic hardware structure diagram of an electronic device provided by an embodiment of the present application. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0018] The following will describe the principles and spirit of the present application 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, not all of the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments herein without creative efforts shall fall within the scope of protection of the present application.

[0019] 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).

[0020] The following briefly describes the relevant content such as the concepts and technical terms that may be involved in the embodiments of this application.

[0021] Orbit mechanics model: A high-precision orbit propagation model established by considering various realistic physical forces (perturbation terms, such as the atmospheric drag perturbation term) acting on a satellite, used to predict the position and velocity of a low-orbit satellite. Common orbit mechanics models include, for example, the Simplified General Perturbation Model 4 (SGP4), the High Precision Orbit Propagator (HPOP), etc.

[0022] Atmospheric drag perturbation: It refers to the drag perturbation generated when a satellite orbits due to its interaction with atmospheric molecules. This perturbation affects the satellite's orbit parameters, such as orbit altitude, orbit inclination, and period. For low-orbit satellites, atmospheric drag is the largest in magnitude among non-conservative forces, so its impact on the satellite's orbit must be considered.

[0023] The small thrust perturbation term refers to the acceleration perturbation generated in the satellite orbit mechanics model due to the continuous low thrust provided by the satellite engine. Different from traditional impulsive thrust, small thrust has the characteristics of high specific impulse, high control precision, and long action time, and is suitable for tasks such as orbit adjustment, attitude control, orbit maintenance, and deep space exploration.

[0024] On-board coordinate system: Also called the UNW (Up-Normal-W) coordinate system, it is a local right-handed coordinate system, and its three axes can be defined as follows: The origin is located at the satellite's center of mass; the x-axis points in the forward direction of the satellite, parallel to the tangent direction of the orbit position where the satellite is located at the corresponding moment; the y-axis points to the right of the satellite's forward direction, perpendicular to the orbit plane of the satellite at the corresponding moment; the z-axis is perpendicular to the x-axis and the y-axis, pointing towards the ground.

[0025] ECI: Earth-Centered Inertial (ECI) coordinate system, which does not rotate with the Earth, has its origin at the Earth's center of mass, and the directions of its three axes can be found.

[0026] Body-fixed Coordinate System (BFC): It is determined relative to the structure and attitude of the satellite itself, moves and rotates with the satellite, and the three axes are defined as follows: The origin is located at the center of mass of the satellite; the x-axis points to the front of the satellite; the y-axis points to the right side of the satellite; the z-axis points to the bottom of the satellite.

[0027] Satellite attitude angles: The attitude of the satellite is described by three angles: yaw angle, roll angle, and pitch angle. Yaw angle: Rotation around the Z-axis, indicating the direction change of the satellite on the horizontal plane. Pitch angle: Rotation around the Y-axis, indicating the deflection angle of the satellite in the vertical direction. Roll angle: Rotation around the X-axis, indicating the rotation of the satellite around the forward direction.

[0028] Feature Fusion: It is the process of combining data features from different sources to enhance the input information of the model. This can be achieved through simple concatenation, weighted combination, or more complex methods (such as dimensionality reduction), aiming to improve the model's understanding and prediction ability for tasks.

[0029] Model fusion technology: Model fusion technology is an ensemble learning method that improves the overall prediction performance by combining the prediction results of multiple models. This method can make up for the deficiencies of a single model and improve the accuracy and robustness of the model.

[0030] Base prediction model: Abbreviated as the base model, each base prediction model is trained and predicted independently of other models, generating its own prediction results based on different learning algorithms or different subsets of data. Different base prediction models have different learning biases and capabilities, and they capture different patterns or features in the data.

[0031] In related technologies, when predicting the position and velocity of a satellite, it usually relies on orbital mechanics models or data-driven orbit prediction methods. Existing orbital mechanics models can be, for example, SGP4 and HPOP. When using orbital mechanics models to predict the position and velocity of a satellite, the following defects exist:

[0032] Only considering atmospheric drag perturbation and conventional perturbation terms (such as the non-spherical gravity of the Earth, the gravity of the Sun and the Moon, etc.), it fails to fully consider the influence of other perturbation factors such as small thrust on the satellite orbit, resulting in a certain deviation between the calculated satellite orbit elements and the actual orbit state, especially in low orbits and high-precision tasks.

[0033] Insufficient adaptability to complex tasks: As the complexity of space missions increases, such as when making orbital adjustments to satellites, attitude control, or performing precise orbital rendezvous missions, the orbital calculation model relying solely on atmospheric drag and conventional perturbation terms may not meet the high-precision requirements.

[0034] Poor adaptability to dynamic environments: During the operation of satellites, the use of small-thrust systems (such as electric propulsion systems) is becoming increasingly common. Existing technologies fail to incorporate small-thrust perturbations into the calculation model, resulting in limitations in the accuracy and reliability of orbital predictions in dynamic orbital adjustments and long-term mission planning.

[0035] In summary, there are problems of insufficient accuracy and poor adaptability in traditional satellite orbital element calculation schemes, especially when facing high-precision missions and complex dynamic environments. Therefore, it is necessary to improve the traditional orbital calculation scheme to enhance the calculation accuracy and reliability of satellite orbital elements.

[0036] Based on this, the present application adds a small-thrust perturbation term to the input parameters of the orbital mechanics model. By combining the small-thrust perturbation term, atmospheric drag perturbation term, and conventional perturbation term, it more comprehensively reflects the actual force conditions of the satellite orbit. This enables the orbital mechanics model to analyze the orbital changes of the satellite under the action of multiple forces according to the actual force conditions, more accurately determine the satellite orbital elements, and make the calculated orbital elements closer to the actual orbital state. Due to the consideration of the small-thrust perturbation term, the calculation of the satellite's position and velocity during the integration process is more accurate, better meeting the high-precision requirements for orbital position calculation in modern space missions to solve at least one of the above technical problems.

[0037] Next, in conjunction with the accompanying drawings, the method for determining the ultra-low orbit of a near-earth satellite provided by the embodiments of the present application will be described in detail through specific embodiments and their application scenarios.

[0038] Figure 1 It is a flowchart showing the method for determining the ultra-low orbit of a near-earth satellite provided by an embodiment of the present application. The execution subject of the method for determining the ultra-low orbit of a near-earth satellite can be a satellite orbit prediction device in an ultra-low orbit guarantee system.

[0039] Next, taking the execution subject of the method for determining the ultra-low orbit of a near-earth satellite as a satellite orbit prediction device as an example, the method for determining the ultra-low orbit of a near-earth satellite of the present application will be described. It should be noted that the above execution subject and application scenario do not limit the present application.

[0040] As Figure 1 shown, the method for determining the ultra-low orbit of a near-earth satellite provided by the embodiments of the present application may include step 110 - step 140.

[0041] Step 110: Obtain the thrust vector of the satellite engine in the satellite body coordinate system when the thrust magnitude of the ion thruster of the satellite is constant.

[0042] Step 120: Obtain the first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and based on the first rotation matrix, transform the thrust vector to the geocentric inertial coordinate system to obtain the thrust components of the thrust vector on the three coordinate axes of the geocentric inertial coordinate system.

[0043] Step 130: Combine the satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes to obtain the small thrust perturbation term.

[0044] Step 140: Add the small thrust perturbation term to the atmospheric drag perturbation term and the conventional drag perturbation term to obtain the input parameters, and input the input parameters into the orbital mechanics model, so that the orbital mechanics model determines the satellite orbit elements based on the atmospheric drag perturbation term, the conventional perturbation term, and the small thrust perturbation term.

[0045] The method for determining the ultra-low orbit of a near-earth satellite provided by the embodiments of the present application takes into account that in the previous satellite orbit position calculation scheme, only the atmospheric drag perturbation term and the conventional drag perturbation term are considered, and the influence of the small thrust perturbation term on the change of the satellite orbit position is not considered. Based on this, the present application adds the small thrust perturbation term to the input parameters of the orbital mechanics model, and combines the small thrust perturbation term, the atmospheric drag perturbation term, and the conventional drag perturbation term to more comprehensively reflect the actual force situation of the satellite orbit, so that the orbital mechanics model can analyze the orbit change of the satellite under the action of multiple forces according to the actual force situation, more accurately determine the satellite orbit elements, and make the calculated orbit elements closer to the actual orbit state. Due to the consideration of the small thrust perturbation term, the calculation of the satellite position and velocity during the integration process is more accurate, better meeting the high-precision requirements of modern space missions for orbit position calculation.

[0046] The following combines specific embodiments to detail the specific implementation manners of the above steps.

[0047] Regarding Step 110: Obtain the thrust vector of the satellite engine in the satellite body coordinate system when the thrust magnitude of the ion thruster of the satellite is constant.

[0048] In Step 110, the thrust generated by the ion thruster is a small thrust. This low thrust is regarded as a perturbation term in orbital dynamics because it causes continuous small perturbations to the orbit of the spacecraft. Although the magnitude of the small thrust is small, it can accumulate a significant velocity increment over a long time. The continuous action of the small thrust will cause the orbit to change slowly. If not considered, it may lead to the accumulation of orbit prediction errors.

[0049] Involving step 120, a first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system is obtained, and based on the first rotation matrix, the thrust vector is converted to the geocentric inertial coordinate system to obtain the thrust components of the thrust vector on the three coordinate axes of the geocentric inertial coordinate system.

[0050] In step 120, the conversion of the same vector between different coordinate systems can be completed by a rotation matrix. The satellite attitude is usually described relative to the ECI coordinate system. Therefore, the present application can obtain the satellite attitude angle through the satellite attitude sensor, and then use the satellite attitude angle to determine the first rotation matrix .

[0051] Specifically, the rotation matrix from three-dimensional coordinate system A to coordinate system B is is a 3×3 matrix that defines how to transform a vector from the coordinate system A to Transformed to coordinate system B The vector can be transformed in the AB coordinate system using formula (1):

[0052] (1)

[0053] Rotation Matrix , in the matrix It represents the projection of the unit vector in the first axis direction of the A coordinate system onto the various coordinate axis directions in the B coordinate system, and so on for the other elements.

[0054] Based on this, the thrust vector of the engine is , the thrust direction of the engine points directly behind the satellite, then The representation in the satellite body coordinate system (BFC) is , T is the thrust generated by the engine.

[0055] Using the first rotation matrix right After coordinate system transformation, we get Representation in the ECI coordinate system , the components of T on the three coordinate axes of ECI can be obtained by calculation:

[0056]

[0057] In step 130, the satellite mass, the real-time on / off status of the engine, and the thrust components on the three coordinate axes are combined to obtain a small thrust perturbation term.

[0058] In step 130, after inputting the small thrust perturbation term into the orbital mechanics model, the orbital mechanics model can calculate the acceleration generated by the small thrust perturbation term based on the satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes.

[0059] Regarding step 140, a small thrust perturbation term is newly added on the basis of the atmospheric drag perturbation term and the conventional drag perturbation term to obtain the input parameters, and the input parameters are input into the orbital mechanics model, so that the orbital mechanics model determines the satellite orbit elements based on the atmospheric drag perturbation term, the conventional perturbation term, and the small thrust perturbation term.

[0060] In step 140, the low-earth orbit satellite orbit transfer has the characteristics of high maneuvering frequency, long continuous maneuvering time, and very small engine thrust. In order to maintain the continuity of orbit prediction, the present application adds a small thrust perturbation term that did not exist in the past to the orbital mechanics model to calculate the acceleration generated by the ion engine thrust, so as to consider the effect of this acceleration on the satellite when determining the satellite orbit elements.

[0061] The conventional perturbation term can include the third-body gravitational perturbation term, the earth's non-spherical perturbation term, the solid tide force perturbation term, the solar radiation pressure perturbation term, and other perturbation terms, etc.

[0062] In some embodiments of the present application, in order to improve the accuracy and reliability of orbit prediction, the above step 140 determines the satellite orbit elements based on the atmospheric drag perturbation term, the conventional perturbation term, and the small thrust perturbation term, and may include Figure 2 Steps 210 - 230 shown.

[0063] Step 210, based on the satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes, determine the first acceleration corresponding to the small thrust perturbation term;

[0064] Step 220, based on the first acceleration corresponding to the small thrust perturbation term and the second acceleration corresponding to the atmospheric drag perturbation term, determine the acceleration generated by the non-conservative force;

[0065] Step 230, based on the acceleration generated by the non-conservative force, the earth's gravitational acceleration, and the third-body gravitational acceleration, determine the satellite orbit elements.

[0066] Specifically, the transmitter can generate a small thrust when it is in the on state, and the engine cannot generate a small thrust when it is in the off state. Therefore, the determined first acceleration is different under different on / off states. The earth's gravitational acceleration and the third-body gravitational acceleration are calculated from the conventional perturbation term.

[0067] In addition to the small thrust perturbation term and the atmospheric drag perturbation term, other non-conservative force perturbation terms can also be included in the non-conservative force. Therefore, the acceleration generated by the non-conservative force can be obtained by adding the first acceleration, the second acceleration, and the acceleration generated by other non-conservative force perturbation terms.

[0068] Exemplarily, if the orbital mechanics model is HPOP, the dynamic equation of HPOP can be described by Newton's second law of formula (2):

[0069] (2)

[0070] Wherein, represents the second derivative of the position vector with respect to time, represents the gravitational acceleration of the earth, , represents the gravitational acceleration of the third celestial body. The atmospheric drag and the small thrust generated by the engine both belong to non-conservative forces, and the accelerations generated by these two forces are represented by and respectively.

[0071] In the embodiments of the present application, a small thrust perturbation term is added on the basis of the atmospheric drag perturbation term and the conventional perturbation term. Based on this, when calculating the satellite orbital elements in the subsequent orbital mechanics model, the first acceleration corresponding to the small thrust perturbation term and the second acceleration corresponding to the atmospheric drag perturbation term can be determined, and then the acceleration generated by the non-conservative force can be determined by combining the first acceleration and the second acceleration. Considering the small thrust perturbation term that was previously ignored, the calculated acceleration generated by the non-conservative force will be more accurate. By comprehensively considering these perturbation factors in the model, the dynamic changes of the satellite orbit can be analyzed more comprehensively, thereby improving the accuracy and reliability of orbit prediction.

[0072] In some embodiments of the present application, in order to improve the accuracy of the first acceleration, step 210 determines the first acceleration corresponding to the small thrust perturbation term based on the satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes, which may specifically include the following steps:

[0073] Substitute the real-time on / off state of the engine into the working state function to obtain a function value, where the on state and the off state correspond to different function values;

[0074] Based on the function value, the satellite mass, and the thrust components on the three coordinate axes, determine the first acceleration corresponding to the small thrust perturbation term.

[0075] Specifically, the acceleration expression of the small thrust perturbation term can be shown as formula (3):

[0076] (3)

[0077] Among them, is the first acceleration, I represents the working state function of the engine, I = 1 when the engine is turned on, and I = 0 when the engine is turned off. The small thrust perturbation term is externally connected to the satellite simulation module to obtain the satellite mass m, the real-time on / off state of the satellite engine, and the thrust magnitude in each direction, so as to calculate .

[0078] In the embodiment of the present application, the thrust magnitude of the ion thruster is constant, so the thrust vector of the engine remains constant. Based on this, the present application can effectively and accurately calculate the first acceleration generated by the small thrust perturbation term by using the thrust vector of the engine, the satellite mass, and the real-time on / off state of the satellite engine.

[0079] In the related art, the traditional method for calculating the windward area usually assumes that the windward area is a fixed value and ignores the influence of satellite attitude changes on the windward area. Therefore, in the previous atmospheric drag perturbation term, the windward area is a fixed constant. However, when the satellite is in orbit, the changes in its attitude, orbital altitude, and the movement of components such as solar panels will cause changes in the windward area. The influence of this change on the atmospheric drag cannot be ignored, especially in the precise orbit determination and orbit prediction of low-orbit satellites. Therefore, if a fixed constant is used as the windward area to calculate the satellite orbit elements, it will affect its accuracy.

[0080] In some embodiments of the present application, in order to further improve the accuracy of the satellite orbit elements, before obtaining the input parameters in step 140 above, the method may further include Figure 3 the steps 310 - 340 shown.

[0081] Step 310, when it is determined that the velocity direction of the satellite is the windward direction, obtain the wind direction unit vector in the satellite body coordinate system, the satellite solar panel area, and the unit normal vector of the plane where the satellite solar panel is located in the satellite body coordinate system;

[0082] Step 320, convert the wind direction unit vector in the satellite body coordinate system to the satellite body coordinate system;

[0083] Step 330, based on the satellite solar panel area, and the wind direction unit vector and unit normal vector in the satellite body coordinate system, determine the real-time windward area of the satellite;

[0084] Step 340, update the fixed windward area in the atmospheric drag perturbation term to the real-time windward area.

[0085] Specifically, if the windward direction is the velocity direction of the satellite, then the wind direction unit vector in the satellite body coordinate system , on this basis, the wind direction unit vector can be converted from the satellite body coordinate system to the satellite body coordinate system by using the rotation matrix.

[0086] In the embodiments of the present application, based on the area of the satellite solar panel, as well as the wind direction unit vector and the unit normal vector in the satellite body coordinate system, the real-time windward area of the satellite is determined, and the fixed windward area in the atmospheric drag perturbation term is updated to the real-time windward area. Compared with a fixed constant, this real-time windward area can be dynamically calculated when the windward area changes due to the movement of components such as the attitude, orbital altitude, and solar panels. Therefore, the accuracy of this real-time windward area is higher. In this way, the present application optimizes and upgrades the atmospheric drag perturbation term. When determining the satellite orbit elements, the acceleration generated by the atmospheric drag perturbation term can be determined based on the more accurate real-time windward area, thereby improving the accuracy of satellite orbit element calculation.

[0087] In some embodiments of the present application, the above step 320 of converting the wind direction unit vector of the on-board coordinate system to the satellite body coordinate system may specifically include:

[0088] Obtain the second rotation matrix of the satellite attitude angle relative to the satellite body coordinate system;

[0089] Based on the second rotation matrix, convert the wind direction unit vector from the on-board coordinate system to the satellite body coordinate system.

[0090] Specifically, the satellite attitude angle includes the roll angle , pitch angle , yaw angle . The second rotation matrix from the on-board coordinate system to the satellite body coordinate system can be calculated by method (4):

[0091] (4)

[0092] where respectively represent the rotation matrices about the z-axis, y-axis, and x-axis of the on-board coordinate system, and are respectively:

[0093]

[0094] Convert the wind direction unit vector from the on-board coordinate system to the satellite body coordinate system, , and through calculation, it can be obtained that .

[0095] In this way, after obtaining the second rotation matrix, the wind direction unit vector can be successfully converted from the on-board coordinate system to the satellite body coordinate system by using the second rotation matrix, and an effective expression of the wind direction unit vector in the satellite body coordinate system can be obtained.

[0096] In some embodiments of the present application, in order to accurately calculate the real-time windward area of the satellite, step 330 above determines the real-time windward area of the satellite based on the satellite solar panel area, as well as the wind direction unit vector and unit normal vector in the satellite body coordinate system, and may specifically include:

[0097] Calculate the absolute value of the product of the wind direction unit vector and the unit normal vector in the satellite body coordinate system;

[0098] Multiply the satellite solar panel area by the absolute value to obtain the real-time windward area of the satellite.

[0099] Specifically, the windward area of the satellite can be calculated using formula (5) :

[0100] (5)

[0101] where S is the solar panel area, and the unit normal vector of the plane where the satellite solar panel is located in the satellite body coordinate system (the general satellite configuration is [0,0,1]), the satellite body can be approximately regarded as a sphere, s is the area of the satellite sphere, and finally the real-time windward area of the satellite is obtained through calculation .

[0102] In the embodiments of the present application, by calculating the absolute value of the product of the wind direction unit vector and the unit normal vector in the satellite body coordinate system and multiplying the satellite solar panel area by the absolute value, the real-time windward area of the satellite can be effectively calculated. In this way, in the scenario where the satellite's attitude changes due to factors such as orbit adjustment and solar panel rotation, by calculating the windward area in real time, these complex attitude changes can be effectively adapted, and the calculation error of the windward area caused by attitude changes can be avoided.

[0103] In some embodiments of the present application, step 140 above determines the satellite orbit elements based on the atmospheric drag perturbation term, the conventional perturbation term, and the small thrust perturbation term, and may specifically include:

[0104] Based on the satellite mass and the real-time windward area in the atmospheric drag perturbation term, determine the real-time windward area-to-mass ratio of the satellite;

[0105] Based on the real-time windward area-to-mass ratio of the satellite, the drag coefficient, the atmospheric density, as well as the speed scalar and speed vector of the satellite relative to the atmosphere, determine the second acceleration corresponding to the atmospheric drag perturbation term.

[0106] Specifically, the acceleration expression of the atmospheric drag perturbation term can be shown as formula (6):

[0107] (6)

[0108] where, is the second acceleration, C represents the drag coefficient, A represents the real-time windward area of the satellite, m represents the mass of the satellite, represents the atmospheric density, v and represent the scalar and vector of the satellite's velocity relative to the atmosphere.

[0109] In the embodiments of the present application, based on the mass and real-time windward area of the satellite in the atmospheric drag perturbation term, the real-time windward area mass ratio of the satellite can be determined, which can more accurately reflect the change of the windward area of the satellite in different attitudes. Then, based on the real-time windward area mass ratio, the second acceleration corresponding to the atmospheric drag perturbation term is determined, which can improve the calculation accuracy of the second acceleration in the satellite orbit mechanics model, and further improve the accuracy of the satellite orbit elements based on the second acceleration.

[0110] In some embodiments of the present application, in order to improve the prediction accuracy of the atmospheric density, before determining the second acceleration corresponding to the atmospheric drag perturbation term based on the real-time windward area mass ratio, drag coefficient, atmospheric density, and the scalar and vector of the satellite's velocity relative to the atmosphere of the satellite, the following steps may further be included: predicting the atmospheric density of the target prediction period by using M basic prediction models to obtain M predicted values, where the M predicted values are output by the M basic prediction models based on the historical environmental parameters corresponding to the target prediction period; assigning M first weight values to the M target predicted values by using a first meta-model; and determining the atmospheric density of the target prediction period by weighted summing the M predicted values by using the M first weight values.

[0111] In the above embodiments, the first meta-model can be trained based on the following steps: obtaining target training data; training the first meta-model by using the target training data, so that the first meta-model assigns corresponding M first weight values to the M predicted values of each prediction period based on the historical environmental parameters corresponding to each prediction period; and determining that the training of the first meta-model is completed when the first loss function value of the first meta-model satisfies a first preset condition, where the first loss function value is the mean square error between the first predicted value and the measured value corresponding to each prediction period in multiple prediction periods, and the first predicted value is obtained by weighted summing the M predicted values corresponding to each prediction period by using the M first weight values assigned by the first meta-model in each prediction period.

[0112] Among them, the target training data includes the historical environmental parameters, measured values, and M predicted values corresponding to each prediction period in multiple prediction periods. The measured value is the atmospheric density actually detected in each prediction period, and the M predicted values are the atmospheric densities predicted by the M basic prediction models for each prediction period based on the historical environmental parameters of each prediction period.

[0113] The first meta-model includes a weight model and a fusion layer. The network structure of the weight model is a Transformer encoder or an LSTM. The weight model is used to assign corresponding M first weight values to M prediction values for different prediction periods. The fusion layer is used to perform weighted summation on the M prediction values by using the M first weight values.

[0114] Corresponding to the method embodiment of the present application, the present application further provides a device for determining the ultra-low orbit of a near-earth satellite.

[0115] Figure 4 It is a schematic structural diagram of a device for determining the ultra-low orbit of a near-earth satellite provided by an embodiment of the present application. As Figure 4 shown, the device 400 for determining the ultra-low orbit of a near-earth satellite may include: an acquisition module 410, a coordinate system conversion module 420, a combination module 430, and a determination module 440.

[0116] Among them, the acquisition module 410 is used to acquire the thrust vector of the satellite engine in the satellite body coordinate system when the thrust magnitude of the ion thruster of the satellite is constant; the coordinate system conversion module 420 is used to acquire the first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and based on the first rotation matrix, convert the thrust vector to the geocentric inertial coordinate system to obtain the thrust components of the thrust vector on the three coordinate axes of the geocentric inertial coordinate system; the combination module 430 is used to combine the satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes to obtain a small thrust perturbation term; the determination module 440 is used to add the small thrust perturbation term on the basis of the atmospheric drag perturbation term and the conventional drag perturbation term to obtain input parameters, and input the input parameters into the orbital mechanics model, so that the orbital mechanics model determines the satellite orbit elements based on the atmospheric drag perturbation term, the conventional perturbation term, and the small thrust perturbation term.

[0117] The device for determining the ultra-low orbit of a near-earth satellite provided by the embodiment of the present application takes into account that in the previous satellite orbit position calculation scheme, only the atmospheric drag perturbation term and the conventional perturbation term are considered, and the influence of the small thrust perturbation term on the change of the satellite orbit position is not considered. Based on this, the present application adds a small thrust perturbation term to the input parameters of the orbital mechanics model. Combining the small thrust perturbation term, the atmospheric drag perturbation term, and the conventional perturbation term can more comprehensively reflect the actual force situation of the satellite orbit, so that the orbital mechanics model can analyze the orbit change of the satellite under the action of various forces according to the actual force situation, and more accurately determine the satellite orbit elements, making the calculated orbit elements closer to the actual orbit state. Due to the consideration of the small thrust perturbation term, the calculation of the satellite position and velocity during the integration process is more accurate, better meeting the high-precision requirements of modern space missions for orbit position calculation.

[0118] In some embodiments of the present application, the determination module includes: a first determination unit configured to determine a first acceleration corresponding to a small thrust perturbation term based on the satellite mass, the real-time on / off state of the engine, and the thrust components on three coordinate axes; a second determination unit configured to determine the acceleration generated by non-conservative forces based on the first acceleration corresponding to the small thrust perturbation term and a second acceleration corresponding to an atmospheric drag perturbation term; and a third determination unit configured to determine the satellite orbit elements based on the acceleration generated by non-conservative forces, the Earth gravitational acceleration, and the third-body gravitational acceleration, wherein the Earth gravitational acceleration and the third-body gravitational acceleration are calculated from conventional perturbation terms.

[0119] In some embodiments of the present application, the first determination unit is specifically configured to: substitute the real-time on / off state of the engine into the working state function to obtain a function value, where different function values correspond to the on state and the off state; and determine the first acceleration corresponding to the small thrust perturbation term based on the function value, the satellite mass, and the thrust components on three coordinate axes.

[0120] Some embodiments of the present application further include: an acquisition module configured to, before obtaining the input parameters, when it is determined that the velocity direction of the satellite is the windward direction, acquire the wind direction unit vector in the on-board coordinate system of the satellite, the satellite solar panel area, and the unit normal vector of the plane where the satellite solar panel is located in the satellite body coordinate system; a coordinate system conversion module further configured to convert the wind direction unit vector in the on-board coordinate system to the satellite body coordinate system; a determination module further configured to determine the real-time windward area of the satellite based on the satellite solar panel area, and the wind direction unit vector and the unit normal vector in the satellite body coordinate system; and an update module configured to update the fixed windward area in the atmospheric drag perturbation term to the real-time windward area.

[0121] In some embodiments of the present application, the coordinate system conversion module is specifically configured to: acquire the attitude strategy of the satellite and determine a second rotation matrix based on the attitude strategy; and convert the wind direction unit vector from the on-board coordinate system to the satellite body coordinate system based on the second rotation matrix.

[0122] In some embodiments of the present application, the determination module is specifically configured to: calculate the absolute value of the product of the wind direction unit vector and the unit normal vector in the satellite body coordinate system; and multiply the satellite solar panel area by the absolute value to obtain the real-time windward area of the satellite.

[0123] In some embodiments of the present application, the determination module is specifically configured to: determine the real-time windward area-to-mass ratio of the satellite based on the satellite mass and the real-time windward area in the atmospheric drag perturbation term; and determine the second acceleration corresponding to the atmospheric drag perturbation term based on the real-time windward area-to-mass ratio of the satellite, the drag coefficient, the atmospheric density, and the speed scalar and speed vector of the satellite relative to the atmosphere.

[0124] The ultra-low orbit determination device for a near-earth satellite provided by an embodiment of the present application can implement Figures 1-3 each process implemented by the service platform in the method embodiment, and can achieve the same technical effects. To avoid repetition, details are not described herein again.

[0125] Figure 5 FIG. is a schematic hardware structure diagram of an electronic device provided by an embodiment of the present application.

[0126] As Figure 5 shown, the electronic device 500 includes a memory 501, a processor 502, and a computer program stored on the memory 501 and executable on the processor 502.

[0127] In one example, the above-mentioned processor 502 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.

[0128] The memory 501 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. Therefore, 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.

[0129] The processor 502 runs a computer program corresponding to the executable program code by reading the executable program code stored in the memory 501, so as to implement the method in the embodiment of the first aspect above.

[0130] In some examples, the electronic device 500 may further include a communication interface 503 and a bus 510. Among them, as Figure 5 shown, the memory 501, the processor 502, and the communication interface 503 are connected through the bus 510 and complete communication with each other.

[0131] The communication interface 503 is mainly used to implement communication between each module, device, unit, and / or device in the embodiments of the present application. The input device and / or output device may also be accessed through the communication interface 503.

[0132] The bus 510 includes hardware, software, or both, and couples the components of the electronic device 500 to each other. By way of example and not limitation, the bus 510 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 510 may include one or more buses. Although the embodiments of the present application describe and illustrate specific buses, the present application contemplates any suitable bus or interconnect.

[0133] The electronic device provided by the embodiments of the present application can implement Figures 1-3 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.

[0134] In combination with the method for determining the ultra-low orbit of a near-earth satellite in the above embodiments, the embodiments of the present application can be implemented by providing a computer storage medium. 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 methods for determining the ultra-low orbit of a near-earth satellite in the above embodiments are implemented.

[0135] In combination with the method for determining the ultra-low orbit of a near-earth satellite in the above embodiments, the embodiments of the present application can be implemented by providing a computer program product. 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 methods for determining the ultra-low orbit of a near-earth satellite in the above embodiments are implemented.

[0136] Another embodiment of the present application further provides a chip, which includes a processor and a communication interface. The communication interface is coupled to the processor. The processor is configured to run programs or instructions to implement each process of the above-mentioned embodiment of the method for determining the ultra-low orbit of a near-earth satellite, and can achieve the same technical effects. To avoid repetition, it will not be elaborated here.

[0137] It should be understood that the chip mentioned in the embodiment of the present application may also be referred to as a system-on-chip, system chip, chip system, or system-on-a-chip, etc.

[0138] 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 between steps after understanding the spirit of the present application.

[0139] It should also be noted that the functional blocks shown in the above 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, 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 via a data signal carried in a carrier wave on a transmission medium or a communication link. A "machine-readable medium" can include any medium capable of storing or transmitting 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, etc. The code segment can be downloaded via a computer network such as the Internet, an intranet, etc.

[0140] 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.

[0141] As 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 combination 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 device to produce a machine, such that the 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 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 is also understood that each block in the block diagram and / or flowchart, and the combination of blocks in the block diagram and / or flowchart, can also be implemented by dedicated hardware that performs the specified functions or actions, or can be implemented by a combination of dedicated hardware and computer instructions.

[0142] As described above, this is only the 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 within the technical scope disclosed in the present application can easily think of various equivalent modifications or substitutions, and these modifications or substitutions should all be covered within the protection scope of the present application.

Claims

1. A method for determining an ultra-low orbit of a near-Earth satellite, characterized in that: include: When the thrust of the satellite's ion thruster is constant, the thrust vector of the satellite engine in the satellite's body coordinate system is obtained; Acquire a first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and transform the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain thrust components of the thrust vector on three coordinate axes of the geocentric inertial coordinate system; The satellite mass, the real-time on / off state of the engine, and the thrust components on the three coordinate axes are combined to obtain a small thrust perturbation term; Adding a small thrust perturbation term on the basis of the atmospheric drag perturbation term and the conventional drag perturbation term to obtain input parameters, and inputting the input parameters into the orbital mechanics model so that the orbital mechanics model determines the satellite orbit elements based on the atmospheric drag perturbation term, the conventional perturbation term and the small thrust perturbation term; Based on the atmospheric drag perturbation, conventional perturbation and small thrust perturbation, the satellite orbit elements are determined, including: Determining a first acceleration corresponding to a small thrust perturbation term based on the mass of the satellite, the real-time on / off state of the engine, and the thrust components on the three coordinate axes; Determining the acceleration generated by the non-conservative force based on the first acceleration corresponding to the small thrust perturbation term and the second acceleration corresponding to the atmospheric drag perturbation term; The satellite orbit elements are determined based on the acceleration generated by the non-conservative force, the earth's gravitational acceleration and the third celestial body's gravitational acceleration, wherein the earth's gravitational acceleration and the third celestial body's gravitational acceleration are calculated by conventional perturbation terms.

2. The method according to claim 1, characterized in that Based on the mass of the satellite, the real-time on / off state of the engine and the thrust components on the three coordinate axes, determining a first acceleration corresponding to the small thrust perturbation term includes: Substitute the real-time on / off state of the engine into the working state function to obtain the function value, wherein the on state and the off state correspond to different function values; A first acceleration corresponding to the small thrust perturbation term is determined based on the function value, the satellite mass and the thrust components on the three coordinate axes.

3. The method according to claim 1, characterized in that Before getting the input parameters, it also includes: When it is determined that the speed direction of the satellite is the windward direction, the wind direction unit vector of the satellite in the on-board coordinate system, the area of ​​the satellite solar panel, and the unit normal vector of the plane where the satellite solar panel is located in the satellite body coordinate system are obtained; Converting the wind direction unit vector of the on-satellite coordinate system to the satellite body coordinate system; Determine the real-time windward area of ​​the satellite based on the area of ​​the satellite solar panel and the wind direction unit vector and unit normal vector in the satellite body coordinate system; The fixed windward area in the atmospheric resistance perturbation term is updated to the real-time windward area.

4. The method according to claim 3, characterized in that The wind direction unit vector of the on-board coordinate system is converted to the satellite body coordinate system, including: Acquire the attitude strategy of the satellite, and determine the second rotation matrix based on the attitude strategy; The wind direction unit vector is transformed from the on-board coordinate system to the satellite body coordinate system based on the second rotation matrix.

5. The method according to claim 3, characterized in that: Based on the satellite solar panel area, and the wind direction unit vector and unit normal vector on the satellite body coordinate system, the real-time windward area of ​​the satellite is determined, including: Calculate the absolute value of the product of the wind direction unit vector and the unit normal vector in the satellite body coordinate system; The satellite solar panel area is multiplied by the absolute value to obtain the real-time windward area of ​​the satellite.

6. The method according to claim 3, characterized in that Based on the atmospheric drag perturbation, conventional perturbation and small thrust perturbation, the satellite orbit elements are determined, including: Based on the satellite mass and the real-time windward area in the atmospheric drag perturbation term, determine the real-time windward surface mass ratio of the satellite; Based on the real-time windward surface mass ratio, drag coefficient, atmospheric density of the satellite, and the velocity scalar and velocity vector of the satellite relative to the atmosphere, a second acceleration corresponding to the atmospheric drag perturbation term is determined.

7. 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 6 is implemented.

8. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores computer program instructions, and when the computer program instructions are executed by a processor, the method according to any one of claims 1 to 6 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 6 when executed.

Images