Ultra-low orbit determination method and device for near-earth satellite, medium and product

Abstract

The invention discloses an ultra-low orbit determination method and device for a near-earth satellite, a medium and a product. Comprising the following steps: acquiring a thrust vector of a satellite engine under a satellite body coordinate system under the condition that the thrust of an ion propeller of a satellite is constant; obtaining a 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 thrust components of the thrust vector on 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 item; and adding a small thrust perturbation item on the basis of the atmospheric resistance perturbation item and the conventional resistance perturbation item to obtain an input parameter, and inputting the input parameter to the orbit mechanical model, so that the orbit mechanical model determines the satellite orbit element based on the atmospheric resistance perturbation item, the conventional perturbation item and the small thrust perturbation item.

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 a method, device, medium and product for determining an ultra-low orbit of a near-Earth satellite. Background Art

[0002] Ultra-low-orbit satellites operate in orbits between 150km and 300km above the Earth's surface. Due to the large number of satellites and the increasing impact of the atmospheric environment, ultra-low-orbit satellites are now facing 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 support system to ensure the stable operation of satellites and efficient mission execution.

[0003] At present, when predicting the future orbital state of a satellite, the existing orbital calculation method mainly relies on atmospheric drag perturbations and conventional perturbations. This calculation method has obvious limitations when faced with high-precision requirements, resulting in a certain deviation between the calculated satellite orbital elements and the actual orbital state, especially in low-orbit and high-precision missions, and cannot meet the high-precision requirements of the satellite orbital position. Summary of the invention

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

[0005] In a first aspect, an embodiment of the present application provides a method for determining an ultra-low orbit of a near-Earth satellite, the method comprising: obtaining a thrust vector of a satellite engine in a satellite body coordinate system when the thrust of an ion thruster of the satellite is constant; obtaining a first rotation matrix of the satellite attitude angle relative to a geocentric inertial coordinate system, and converting the thrust vector to a 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; combining the satellite mass, the real-time on / off status of the engine, and the thrust components on the three coordinate axes to obtain a small thrust perturbation term; adding a small thrust perturbation term on the basis of an atmospheric drag perturbation term and a 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 number of satellite orbital elements based on the atmospheric drag perturbation term, the conventional perturbation term, and the small thrust perturbation term.

[0006] In a second aspect, an embodiment of the present application provides an ultra-low orbit determination device for a near-Earth satellite, the device comprising: an acquisition module, used to acquire the thrust vector of a satellite engine in a satellite body coordinate system when the thrust of an ion thruster of the satellite is constant; a coordinate system conversion module, used to acquire a first rotation matrix of the satellite attitude angle relative to a geocentric inertial coordinate system, and convert the thrust vector to a 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; a combination module, used to combine the satellite mass, the real-time on / off status of the engine, and the thrust components on the three coordinate axes to obtain a small thrust perturbation item; a determination module, used to add a small thrust perturbation item on the basis of an atmospheric drag perturbation item and a conventional drag perturbation item to obtain input parameters, and input the input parameters to an orbital mechanics model, so that the orbital mechanics model determines the number of satellite orbital elements based on the atmospheric drag perturbation item, the conventional perturbation item, and the small thrust perturbation item.

[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 method for determining an ultra-low orbit of a near-Earth 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 having computer program instructions stored thereon. When the computer program instructions are executed by a processor, the steps of the method for determining an ultra-low orbit of a near-Earth 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 method for determining an ultra-low orbit of a near-Earth 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 method for determining an ultra-low orbit of a near-Earth satellite as in the first aspect.

[0011] The present application provides a method, device, equipment, medium and product for determining the ultra-low orbit of a near-Earth satellite, taking into account that in the previous satellite orbit position calculation scheme, only atmospheric drag perturbation terms and conventional perturbation terms were considered, and the influence of small thrust perturbation terms on the change of satellite orbit position was not considered. Based on this, the present application adds a 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 perturbation term to more comprehensively reflect the actual force conditions of the satellite orbit, so that the orbital mechanics model can analyze the orbital changes of the satellite under the action of various forces according to the actual force conditions, and more accurately determine the satellite orbital elements, so that the calculated orbital elements are closer to the actual orbital state. Since the small thrust perturbation term is taken into account, the calculation of the satellite position and velocity in the integration process is more accurate, which better meets the high-precision requirements of modern aerospace missions for orbital position calculation. 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 a method for determining an ultra-low orbit of a near-Earth satellite provided in one embodiment of the present application; Figure 2 is a flow chart of a method for determining an ultra-low orbit of a near-Earth satellite provided in another embodiment of the present application; Figure 3 is a flowchart of a method for determining an ultra-low orbit of a near-Earth satellite provided in yet another embodiment of the present application; Figure 4 It is a structural schematic diagram of an ultra-low orbit determination device for a near-Earth satellite provided in an embodiment of the present application; Figure 5 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] Orbital mechanics model: A high-precision orbit propagation model is established by considering the various real physical forces (perturbations, such as atmospheric drag perturbations) that the satellite is subject to. It is used to predict the position and velocity of low-orbit satellites. Common orbital mechanics models include the Simplified General Perturbation Model 4 (SGP4), the High Precision Orbit Propagator (HPOP), etc.

[0018] Atmospheric drag perturbation: refers to the drag perturbation caused by the interaction between the satellite and the atmospheric molecules when the satellite is orbiting. This perturbation will affect the satellite's orbital parameters, such as orbital altitude, orbital inclination and period. For low-orbit satellites, atmospheric drag is the largest non-conservative force, so its impact on the satellite orbit must be considered.

[0019] The low-thrust perturbation term refers to the acceleration perturbation caused by the continuous low thrust provided by the satellite engine in the satellite orbital mechanics model. Different from the traditional pulse thrust, the low-thrust thrust has the characteristics of high specific impulse, high control accuracy and long action time, and is suitable for orbit adjustment, attitude control, orbit maintenance and deep space exploration.

[0020] On-board coordinate system: also called UNW (Up-Normal-W) coordinate system, is a local right-handed coordinate system whose three axes can be defined as follows: the origin is located at the satellite's center of mass; the x-axis points in the direction of the satellite's forward motion, parallel to the tangent direction of the satellite's orbital position at the corresponding moment; the y-axis points to the right of the satellite's forward motion direction, perpendicular to the satellite's orbital plane at the corresponding moment; the z-axis is perpendicular to the x-axis and y-axis and points to the ground.

[0021] ECI: Earth-Centered Inertial (ECI) coordinate system, which does not rotate with the Earth. The origin is located at the center of mass of the Earth, and the directions of the three axes can be checked.

[0022] Satellite body coordinate system (BFC): determined relative to the satellite's own structure and attitude, moves and rotates with the satellite, and the three axes are defined as follows: the origin is located at the satellite's center of mass; the x-axis points directly in front of the satellite; the y-axis points to the right of the satellite; and the z-axis points directly below the satellite.

[0023] Satellite attitude angle: The attitude of the satellite is described by three angles: yaw angle, roll angle and pitch angle. Yaw angle (Yawangle): Rotation around the Z axis, indicating the change in the direction of the satellite in the horizontal plane. Pitch angle (Pitchangle): Rotation around the Y axis, indicating the deflection angle of the satellite in the vertical direction. Roll angle (Rollangle): Rotation around the X axis, indicating the rotation of the satellite around the forward direction.

[0024] 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 of the task.

[0025] Model fusion technology: Model fusion technology is an integrated learning method that improves the overall prediction performance by combining the prediction results of multiple models. This method can make up for the shortcomings of a single model and improve the accuracy and robustness of the model.

[0026] Basic prediction model: referred to as base model, each basic prediction model is trained and predicted independently of other models, and generates its own prediction results based on different learning algorithms or different data subsets. Different basic prediction models have different learning biases and capabilities, and they capture different patterns or features in the data.

[0027] In the related art, when predicting the position and velocity of a satellite, it is usually based on an orbital mechanics model or a data-driven orbit prediction method. Existing orbital mechanics models, such as SGP4 and HPOP, have the following defects when predicting the position and velocity of a satellite using orbital mechanics models: Only considering atmospheric drag perturbations and conventional perturbations (such as the non-spherical gravity of the Earth, the gravity of the Sun and the Moon, etc.), but failing to fully consider the impact of other perturbations such as small thrust on the satellite orbit, results in a certain deviation between the calculated satellite orbit elements and the actual orbital state, especially in low-orbit and high-precision missions.

[0028] Insufficient adaptability to complex missions: As the complexity of space missions increases, such as orbit adjustment, attitude control, or precise orbit rendezvous of satellites, orbit calculation models that rely solely on atmospheric drag and conventional perturbations may not be able to meet high-precision requirements.

[0029] Poor adaptability to dynamic environments: During satellite operations, the use of small thrust systems (such as electric propulsion systems) is becoming more and more common. Existing technologies fail to incorporate small thrust perturbations into the calculation model, resulting in limited accuracy and reliability of orbit prediction in dynamic orbit adjustments and long-term mission planning.

[0030] In summary, the traditional satellite orbit element calculation scheme has the problems of insufficient accuracy and poor adaptability, especially when facing high-precision tasks and complex dynamic environments. Therefore, it is necessary to improve the traditional orbit calculation scheme to improve the calculation accuracy and reliability of satellite orbit elements.

[0031] Based on this, the present application adds a 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 perturbation term to more comprehensively reflect the actual force conditions of the satellite orbit, so that the orbital mechanics model can analyze the orbital changes of the satellite under the action of various forces according to the actual force conditions, and more accurately determine the satellite orbital elements, so that the calculated orbital elements are closer to the actual orbital state. Due to the consideration of the small thrust perturbation term, the calculation of the satellite position and velocity in the integration process is more accurate, which better meets the high-precision requirements of modern aerospace missions for orbital position calculation, so as to solve at least one of the above technical problems.

[0032] The following, in conjunction with the accompanying drawings, describes in detail the method for determining the ultra-low orbit of a near-Earth satellite provided in the embodiment of the present application through specific embodiments and their application scenarios.

[0033] Figure 1 It is a flow chart of a method for determining an ultra-low orbit of a near-Earth satellite provided in one embodiment of the present application. The executor of the method for determining an ultra-low orbit of a near-Earth satellite may be a satellite orbit prediction device in an ultra-low orbit guarantee system.

[0034] The following takes the implementation subject of the method for determining the ultra-low orbit of a near-Earth satellite as a satellite orbit prediction device as an example to illustrate the method for determining the ultra-low orbit of a near-Earth satellite of the present application. It should be noted that the above implementation subject and application scenario do not constitute a limitation on the present application.

[0035] like Figure 1 As shown, the method for determining the ultra-low orbit of a near-Earth satellite provided in an embodiment of the present application may include steps 110 to 140.

[0036] Step 110, 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; Step 120, obtaining a 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 thrust components of the thrust vector on three coordinate axes of the geocentric inertial coordinate system; Step 130, combining the satellite mass, the real-time on / off status of the engine, and the thrust components on the three coordinate axes to obtain a small thrust perturbation term; Step 140, adding a small thrust perturbation term on the basis of the atmospheric drag perturbation term and the conventional drag perturbation term, obtaining 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.

[0037] The ultra-low orbit determination method for a near-Earth satellite provided in an embodiment of the present application takes into account that in previous satellite orbit position calculation schemes, only atmospheric drag perturbations and conventional perturbations were considered, and the influence of small thrust perturbations on the change of satellite orbit position was not considered. Based on this, the present application adds a 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 perturbation term to more comprehensively reflect the actual force conditions of the satellite orbit, so that the orbital mechanics model can analyze the orbital changes of the satellite under the action of various forces according to the actual force conditions, and more accurately determine the satellite orbit elements, so that the calculated orbit elements are closer to the actual orbit state. Since the small thrust perturbation term is taken into account, the calculation of the satellite position and velocity in the integration process is more accurate, which better meets the high-precision requirements of modern aerospace missions for orbital position calculations.

[0038] The specific implementation of the above steps will be described in detail below in conjunction with specific embodiments.

[0039] Involving step 110, when the thrust magnitude of the satellite's ion thruster is constant, the thrust vector of the satellite engine in the satellite's body coordinate system is obtained.

[0040] In step 110, the thrust generated by the ion thruster is a small thrust, which is regarded as a perturbation term in orbital dynamics because it produces a continuous small disturbance to the orbit of the spacecraft. Although the small thrust is small in magnitude, it can accumulate significant velocity increments over a long period of time. The continuous effect of the small thrust will cause a slow change in the orbit, which may lead to the accumulation of orbit prediction errors if not taken into account.

[0041] 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.

[0042] 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 .

[0043] 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): (1) 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.

[0044] 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.

[0045] 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:

[0046] 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.

[0047] In step 130, after the small thrust perturbation term is input 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 status of the engine, and the thrust components on the three coordinate axes.

[0048] Involving step 140, a small thrust perturbation term is added on the basis of the atmospheric drag perturbation term and the conventional drag perturbation term to obtain 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.

[0049] In step 140, the ultra-low orbit satellite orbit change 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 was not previously available 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.

[0050] Conventional perturbations may include third-body gravitational perturbations, the Earth's non-spherical perturbations, solid tidal perturbations, solar radiation pressure perturbations, and other perturbations.

[0051] 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, which may include: Figure 2 Steps 210 to 230 are shown.

[0052] Step 210, determining a 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; Step 220, 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; Step 230, determining the satellite orbital elements based on the acceleration generated by non-conservative forces, the earth's gravitational acceleration and the gravitational acceleration of a third celestial body.

[0053] Specifically, the engine can generate a small thrust when it is in the on state, and cannot generate a small thrust when it is in the off state, so the first acceleration determined is different in different on / off states. The gravitational acceleration of the earth and the gravitational acceleration of the third celestial body are calculated by conventional perturbation terms.

[0054] In addition to the small thrust perturbation term and the atmospheric drag perturbation term, the non-conservative force may also include other non-conservative force perturbation terms. 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.

[0055] For example, the orbital mechanics model is HPOP, and the dynamic equation of HPOP can be described by Newton's second law of motion in formula (2): (2) in, represents the second-order derivative of the position vector with respect to time, is the Earth's gravitational acceleration, , represents the gravitational acceleration of the third celestial body. The atmospheric drag and the small thrust generated by the engine are both non-conservative forces, and are represented by and to represent the acceleration produced by these two forces.

[0056] In the embodiment 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 the satellite orbit elements are calculated by 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 is determined by combining the first acceleration and the second acceleration. The small thrust perturbation term that was previously ignored is taken into account, so 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 the orbit prediction.

[0057] In some embodiments of the present application, in order to improve the accuracy of the first acceleration, the above 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, and may specifically include the following steps: 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; Based on the function value, the satellite mass and the thrust components on the three coordinate axes, the first acceleration corresponding to the small thrust perturbation term is determined.

[0058] Specifically, the acceleration expression of the small thrust perturbation term can be shown as formula (3): (3) in, is the first acceleration, I represents the working state function of the engine, I=1 when it is turned on, and I=0 when it is turned off. The small thrust perturbation term is 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 size in each direction, so as to calculate .

[0059] In an embodiment of the present application, the thrust of the ion thruster is constant, so the thrust vector of the engine is constant. Based on this, the present application can use the thrust vector of the engine, the mass of the satellite, and the real-time on / off status of the satellite engine to effectively and accurately calculate the first acceleration generated by the small thrust perturbation term.

[0060] In related technologies, traditional methods for calculating the windward area usually assume that the windward area is a fixed value, ignoring the impact 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 a satellite is in orbit, its attitude, orbital altitude, and the movement of components such as solar panels will cause changes in the windward area. The impact of this change on 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 orbital elements, its accuracy will be affected.

[0061] 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, the method may further include: Figure 3 Steps 310 to 340 are shown.

[0062] Step 310, when it is determined that the speed direction of the satellite is the windward direction, obtain 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; Step 320, converting the wind direction unit vector of the on-board coordinate system to the satellite body coordinate system; Step 330, determining 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; Step 340: Update the fixed windward area in the atmospheric resistance perturbation term to the real-time windward area.

[0063] Specifically, if the windward direction is the satellite's velocity direction, then the wind direction unit vector in the satellite coordinate system is On this basis, the wind direction unit vector can be transformed from the on-board coordinate system to the satellite body coordinate system using the rotation matrix.

[0064] In the embodiment of the present application, the real-time windward area of ​​the satellite is determined 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, and the fixed windward area in the atmospheric resistance perturbation term is updated to the real-time windward area. Compared with the fixed constant, the real-time windward area can be dynamically calculated when the attitude, orbital altitude and the movement of components such as solar panels cause the windward area to change, so the accuracy of the real-time windward area is higher. In this way, the present application optimizes and upgrades the atmospheric resistance perturbation term. When determining the satellite orbital elements, the acceleration generated by the atmospheric resistance perturbation term can be determined based on the more accurate real-time windward area, thereby improving the accuracy of the calculation of the satellite orbital elements.

[0065] In some embodiments of the present application, the above step 320 converts the wind direction unit vector of the on-board coordinate system to the satellite body coordinate system, which may specifically include: Obtain a second rotation matrix of the satellite attitude angle relative to the satellite body coordinate system; 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.

[0066] 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 It can be calculated by method (4): (4) in, They represent the rotation matrices around the z-axis, y-axis, and x-axis of the on-satellite coordinate system, respectively:

[0067] Convert the wind direction unit vector from the on-board coordinate system to the satellite body coordinate system. , after calculation, we can get .

[0068] 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 using the second rotation matrix, thereby obtaining an effective expression of the wind direction unit vector in the satellite body coordinate system.

[0069] In some embodiments of the present application, in order to accurately calculate the real-time windward area of ​​the satellite, the above step 330 determines 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, which may specifically include: Calculate the absolute value of the product of the wind direction unit vector and the unit normal vector in the satellite body coordinate system; Multiply the satellite solar panel area by the absolute value to obtain the real-time windward area of ​​the satellite.

[0070] Specifically, the satellite's windward area can be calculated using formula (5): : (5) Where S is the area of ​​the solar panel, and the unit normal vector of the plane where the satellite's solar panel is located in the satellite's body coordinate system is (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 the real-time windward area of ​​the satellite is finally obtained by calculation .

[0071] In the embodiment of the present application, the absolute value of the product of the wind direction unit vector and the unit normal vector in the satellite body coordinate system is calculated, and the satellite solar panel area is multiplied by the absolute value, so that the real-time windward area of ​​the satellite can be effectively calculated. In this way, in the scenario where the satellite changes its attitude due to factors such as orbit adjustment and rotation of solar panels, by calculating the windward area in real time, these complex attitude changes can be effectively adapted to avoid the calculation error of the windward area caused by attitude changes.

[0072] In some embodiments of the present application, 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, which may specifically include: 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; The second acceleration corresponding to the atmospheric drag perturbation term is determined based on the satellite's real-time windward surface mass ratio, drag coefficient, atmospheric density, and the satellite's velocity scalar and velocity vector relative to the atmosphere.

[0073] Specifically, the acceleration expression of the atmospheric drag perturbation term can be shown as formula (6): (6) in, is the second acceleration, C is the drag coefficient, A is the real-time frontal area of ​​the satellite, and m is the mass of the satellite. represents the atmospheric density, v and Represents the velocity scalar and velocity vector of the satellite relative to the atmosphere.

[0074] In an embodiment of the present application, based on the satellite mass and real-time windward area in the atmospheric drag perturbation term, the real-time windward surface mass ratio of the satellite can be determined, which can more accurately reflect the change in the windward area of ​​the satellite under different attitudes. Then, based on the real-time windward surface 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 orbital mechanics model, and then improve the accuracy of the satellite orbital elements based on the second acceleration.

[0075] In some embodiments of the present application, in order to improve the forecast accuracy of atmospheric density, before determining the second acceleration corresponding to the atmospheric drag perturbation term based on the satellite's real-time upwind surface mass ratio, drag coefficient, atmospheric density, and the satellite's velocity scalar and velocity vector relative to the atmosphere, the following steps may also be included: using M basic prediction models to predict the atmospheric density of a target prediction period to obtain M prediction values, wherein the M prediction values ​​are output by the M basic prediction models based on historical environmental parameters corresponding to the target prediction period; assigning M first weight values ​​to the M target prediction values ​​through a first meta-model; and using the M first weight values ​​to perform weighted summation on the M prediction values ​​to determine the atmospheric density of the target prediction period.

[0076] In the above embodiment, the first metamodel can be trained based on the following steps: obtaining target training data; using the target training data to train the first metamodel so that the first metamodel assigns corresponding M first weight values ​​to the M prediction values ​​of each prediction period based on the historical environmental parameters corresponding to each prediction period; when the first loss function value of the first metamodel satisfies the first preset condition, it is determined that the training of the first metamodel is completed, and the first loss function value is the mean square error between the first prediction value corresponding to each prediction period in multiple prediction periods and the measured value, and the first prediction value is obtained by weighted summing the M prediction values ​​corresponding to each prediction period using the M first weight values ​​assigned by the first metamodel in each prediction period.

[0077] Among them, the target training data includes historical environmental parameters, measured values ​​and M predicted values ​​corresponding to each prediction period in multiple prediction periods. The measured values ​​are the atmospheric density of each prediction period actually detected, and the M predicted values ​​are obtained by predicting the atmospheric density of each prediction period based on the historical environmental parameters of each prediction period by M basic prediction models.

[0078] The first meta-model includes a weight model and a fusion layer. The network structure of the weight model is transformer encoder or Lstm. The weight model is used to assign corresponding M first weight values ​​to M prediction values ​​of different prediction periods. The fusion layer is used to perform weighted summation on the M prediction values ​​using the M first weight values.

[0079] Corresponding to the method embodiment of the present application, the present application also provides an ultra-low orbit determination device for a near-Earth satellite.

[0080] Figure 4 Schematic diagram of the structure of a device for determining an ultra-low orbit of a near-Earth satellite provided in an embodiment of the present application. Figure 4 As shown, the ultra-low orbit determination device 400 of the near-Earth satellite may include: an acquisition module 410, a coordinate system conversion module 420, a combination module 430, and a determination module 440.

[0081] Among them, the acquisition module 410 is used to obtain the thrust vector of the satellite engine in the satellite body coordinate system when the thrust size of the satellite's ion thruster is constant; the coordinate system conversion module 420 is used to obtain 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; the combination module 430 is used to combine the satellite mass, the real-time on / off status of the engine, and the thrust components on the three coordinate axes to obtain a small thrust perturbation item; the determination module 440 is used to add a small thrust perturbation item on the basis of the atmospheric drag perturbation item and the conventional drag perturbation item to obtain input parameters, and input the input parameters to the orbital mechanics model, so that the orbital mechanics model determines the satellite orbital elements based on the atmospheric drag perturbation item, the conventional perturbation item and the small thrust perturbation item.

[0082] The ultra-low orbit determination device for a near-Earth satellite provided in an embodiment of the present application takes into account that in the previous satellite orbit position calculation scheme, only atmospheric drag perturbation terms and conventional perturbation terms were considered, and the influence of small thrust perturbation terms on the change of satellite orbit position was not considered. Based on this, the present application adds a 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 perturbation term to more comprehensively reflect the actual force conditions of the satellite orbit, so that the orbital mechanics model can analyze the orbital changes of the satellite under the action of various forces according to the actual force conditions, and more accurately determine the satellite orbit elements, so that the calculated orbit elements are closer to the actual orbit state. Since the small thrust perturbation term is taken into account, the calculation of the satellite position and velocity in the integration process is more accurate, which better meets the high-precision requirements of modern aerospace missions for orbital position calculation.

[0083] In some embodiments of the present application, the determination module includes: a first determination unit, used to determine the first acceleration corresponding to the small thrust perturbation term based on the satellite mass, the real-time on / off status of the engine, and the thrust components on the three coordinate axes; a second determination unit, used to determine 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; a third determination unit, used to determine the satellite orbital elements 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.

[0084] In some embodiments of the present application, the first determination unit is specifically used to: substitute the real-time on / off state of the engine into the working state function to obtain a function value, wherein the on state and the off state correspond to different function values; 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.

[0085] Some embodiments of the present application also include: an acquisition module, which is used to, before obtaining the input parameters, obtain 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 when it is determined that the speed direction of the satellite is the windward direction; a coordinate system conversion module, which is also used to convert the wind direction unit vector of the on-board coordinate system to the satellite body coordinate system; a determination module, which is also used to determine the real-time windward area of ​​the satellite based on the satellite solar panel area, and the wind direction unit vector and unit normal vector in the satellite body coordinate system; an update module, which is used to update the fixed windward area in the atmospheric resistance perturbation term to the real-time windward area.

[0086] In some embodiments of the present application, the coordinate system conversion module is specifically used to: obtain the attitude strategy of the satellite, and determine the second rotation matrix based on the attitude strategy; based on the second rotation matrix, convert the wind direction unit vector from the on-board coordinate system to the satellite body coordinate system.

[0087] In some embodiments of the present application, the determination module is specifically used 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; multiply the satellite solar panel area by the absolute value to obtain the real-time windward area of ​​the satellite.

[0088] In some embodiments of the present application, the determination module is specifically used to: determine the real-time windward surface mass ratio of the satellite based on the satellite mass and the real-time windward area in the atmospheric drag perturbation term; determine the second acceleration corresponding to the atmospheric drag perturbation term based on the satellite's real-time windward surface mass ratio, drag coefficient, atmospheric density, and the velocity scalar and velocity vector of the satellite relative to the atmosphere.

[0089] The ultra-low orbit determination device for near-Earth satellites provided in the embodiment of the present application can achieve Figure 1-3 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.

[0090] Figure 5 It is a schematic diagram of the hardware structure of an electronic device provided in an embodiment of the present application.

[0091] like Figure 5 As shown, the electronic device 500 includes a memory 501 , a processor 502 , and a computer program stored in the memory 501 and executable on the processor 502 .

[0092] In an example, the processor 502 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.

[0093] The memory 501 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.

[0094] 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.

[0095] In some examples, the electronic device 500 may further include a communication interface 503 and a bus 510. Figure 5 As shown, the memory 501, the processor 502, and the communication interface 503 are connected via a bus 510 and communicate with each other.

[0096] The communication interface 503 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 503 can also be used to access input devices and / or output devices.

[0097] The bus 510 includes hardware, software, or both, coupling 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 embodiments of the present application describe and illustrate a particular bus, the present application contemplates any suitable bus or interconnect.

[0098] The electronic device provided in the embodiment of the present application can realize Figure 1-3 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.

[0099] In combination with the method for determining the ultra-low orbit of a near-Earth 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 a processor, the steps of any one of the methods for determining the ultra-low orbit of a near-Earth satellite in the above embodiment are implemented.

[0100] In combination with the method for determining the ultra-low orbit of a near-Earth 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 methods for determining the ultra-low orbit of a near-Earth satellite in the above embodiment are implemented.

[0101] 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 embodiment of the method for determining the ultra-low orbit of a near-Earth satellite, and can achieve the same technical effect. To avoid repetition, it will not be repeated here.

[0102] 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.

[0103] 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.

[0104] 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.

[0105] 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.

[0106] 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.

[0107] 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 this application, and these modifications or replacements should be included in the protection scope of this 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; A small thrust perturbation term is added on the basis of the atmospheric drag perturbation term and the conventional drag perturbation term to obtain input parameters, and the input parameters are input 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.

2. The method according to claim 1, characterized in that 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.

3. The method according to claim 2, 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.

4. 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.

5. The method according to claim 4, 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.

6. The method according to claim 4, 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.

7. The method according to claim 4, 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.

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

9. A computer-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 7 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 7 when executed.

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