A method for improving orbit accuracy of low-orbit detectors

Abstract

The present invention discloses a method for improving the orbit accuracy of a low-orbit detector, which belongs to the field of deep space navigation detection; the method comprises: calculating the acceleration generated by each sailboard of the detector, and accumulating them to obtain the acceleration generated by the light pressure of the detector; calculating the acceleration caused by the planetary radiation pressure on each sailboard of the detector and the acceleration accumulation of the albedo and infrared radiation force suffered by the detector to obtain the total acceleration of the detector; performing numerical integration according to the total acceleration of the detector to obtain the calculated value; calculating the centroid correction value from the phase center of the antenna to the centroid of the detector, adding the centroid correction value to the calculated value to obtain the corrected calculated value; using the corrected calculated value to continuously iterate the calculation to obtain a high-precision precise orbit. The present invention uses a high-precision multi-sailboard model to model each part of the detector and perform high-gain antenna correction, which can not only obtain the precise light pressure of the detector, but also perform centroid correction, thereby improving the overall orbit accuracy of the detector.

Description

Technical Field

[0001] The invention relates to a method for improving the orbit accuracy of a low-orbit detector, belonging to the technical field of deep space navigation detection. Background Art

[0002] At present, the precise orbit determination of lunar and Martian probes generally uses a set of force models to numerically integrate the probe's orbit to obtain a calculated value (i.e., the expected measurement value), subtracts it from the observed value obtained by actual observation to obtain the residual, and iterates repeatedly through the least squares method to minimize the sum of squares of the residual, thereby obtaining the final precise orbit. Therefore, the most important thing in precise orbit determination is to use a precise force model, and one of the most influential forces in the force model is solar light pressure. Solar light pressure refers to the thrust exerted on the probe by sunlight shining on it. However, in current research, the light pressure model used is often the most simplified, that is, the total area of ​​the probe is used as the illuminated area. Under this model, there are only two possibilities for the light pressure received by the probe: fully illuminated and not illuminated (i.e., the illuminated area is 0). The calculated light pressure plot is displayed as a straight line (or no), which is obviously not in line with the actual situation. When the probe is moving, it will be partially illuminated. At this time, the light pressure plot it receives should be displayed with fluctuations, so the light pressure calculated by this model will have obvious errors. Therefore, modeling the detector to obtain its precise light pressure is one of the challenges currently faced in the field of precise orbit determination, and it is also a difficult problem in the field of deep space navigation and detection.

[0003] In addition, current research often overlooks a problem in precise orbit determination, that is, the antenna that receives data for measurement is not the same point as the center of the precise orbit determination, so a center of mass correction is required to calculate the correction from the phase center of the antenna to the center of mass of the detector. Since the high-gain antenna is controlled and rotated by multiple robotic arms, this correction cannot be simply added with a fixed constant. How to calculate the center of mass correction value is also a difficult problem that needs to be solved in precise orbit determination. Summary of the invention

[0004] The purpose of the present invention is to address the problems of inaccurate light pressure and center of mass offset in the current precise orbit determination of lunar and Martian probes, and to propose a precise orbit determination method for lunar and Martian probes based on a high-precision multi-sailboard model. This method will use the external attitude data of the probe to establish a high-precision model for the probe, thereby calculating the light pressure and center of mass offset, and finally making corrections to obtain a more precise orbit with higher accuracy.

[0005] In order to achieve the above purpose / solve the above technical problems, the present invention is implemented by adopting the following technical solutions:

[0006] A method for improving the orbit accuracy of a low-orbit detector, the method comprising:

[0007] Based on a high-precision multi-board model, the sum of the pressure generated by the incident sunlight and the pressure generated by solar radiation on each board of the detector is calculated to obtain the acceleration generated by each board.

[0008] Accumulate the acceleration generated by each sailboard to obtain the acceleration generated by the detector light pressure;

[0009] Based on a high-precision multi-panel model, the acceleration of each panel of the probe caused by the planetary radiation pressure is calculated;

[0010] The acceleration of each sailboard caused by the planetary radiation pressure is accumulated to obtain the acceleration of the albedo and infrared radiation force on the probe;

[0011] The total acceleration of the detector is obtained by adding the acceleration generated by the light pressure of the detector and the acceleration of the albedo and infrared radiation force on the detector.

[0012] According to the total acceleration of the detector, numerical integration is performed to obtain the calculated value;

[0013] According to the external attitude data of the detector, the centroid correction value from the phase center of the antenna to the centroid of the detector is calculated, and the centroid correction value is added to the calculated value to obtain the corrected calculated value;

[0014] The corrected calculated values ​​are used for continuous iterative calculations to obtain a high-precision precise orbit.

[0015] Optionally, the acceleration generated by each windsurfing board is calculated as:

[0016]

[0017] Among them, a i represents the acceleration generated by each sailboard, i is the number of the probe sailboard, there are N sailboards in total, C r is the solar radiation coefficient, P s is the solar pressure constant at 1 astronomical unit, A u is the length of an astronomical unit; r s is the distance from the sun to the satellite; A is the area of ​​the sailboard; θ is the angle between the normal vector of the sailboard and the direction of the sun, m is the mass of the satellite, δ is the diffuse reflectivity, ρ is the specular reflectivity, is the normal vector of the sailboard; is the unit vector pointing from the satellite to the sun.

[0018] Optionally, the acceleration generated by each sailboard is accumulated to obtain the calculation formula of the acceleration generated by the detector light pressure:

[0019]

[0020] Among them, a s represents the acceleration generated by the light pressure of the detector, i is the number of the detector sailboard, there are N sailboards in total, C r is the solar radiation coefficient, P s is the solar pressure constant at 1 astronomical unit, A u is the length of an astronomical unit, r s is the distance from the sun to the satellite, θ i is the angle between the normal vector of sailboard i and the direction of the sun, A i is the area of ​​sailboard i, m is the mass of satellite, δ is diffuse reflectivity, ρ is specular reflectivity, is the normal vector of the sailboard i, is the unit vector pointing from the satellite to the sun.

[0021] Optionally, the calculation formula for calculating the acceleration of each sailboard of the detector caused by the planetary radiation pressure based on the high-precision multi-sailboard model is:

[0022]

[0023] in, represents the acceleration of each sailboard caused by the planetary radiation pressure, i is the number of the probe sailboard, there are N sailboards in total, j is the block divided by the planetary surface, there are J in total max surface blocks, G j is the radiation flux of the radiation source, A is the area of ​​the sailboard, m is the mass of the satellite, c is the speed of light, δ is the diffuse reflectivity, ρ is the mirror reflectivity, θ j represents the angle between the normal vector of the sailboard and the vector of the block j pointing to the planet surface, is the normal vector of the sailboard, is the unit vector pointing from the satellite to the source of the radiation.

[0024] Optionally, the acceleration of each sailboard caused by the planetary radiation pressure is accumulated to obtain the acceleration calculation formula of the albedo and infrared radiation force on the detector:

[0025]

[0026] Among them, a a It represents the acceleration of the albedo and infrared radiation force on the probe, i is the number of the probe sailboard, there are N sailboards in total, j is the block divided by the planetary surface, there are J in total max surface blocks, G j is the radiation flux of the radiation source, A i is the area of ​​sailboard i, m is the mass of the satellite, c is the speed of light, δ is the diffuse reflectivity, ρ is the mirror reflectivity, and ρ iis the mirror reflectivity of sailboard i, θ ij represents the angle between the normal vector of sailboard i and the vector of sailboard pointing to block j on the surface of Mars. If θ ij Greater than 90 。 , then the acceleration is 0, is the normal vector of the sailboard, is the unit vector pointing from the satellite to the source of the radiation.

[0027] Optionally, the method of calculating a centroid correction value from the phase center of the antenna to the centroid of the detector based on the high-precision multi-sailboard model and the external attitude data of the detector includes:

[0028] Based on the detector design data, the initial offset vector of the detector is obtained:

[0029]

[0030] Wherein, offset is the initial offset vector, HGA1 corresponds to link1. Since link1 is connected to the detector body, it is directly regarded as the first component of the centroid correction before projection. HGA2 and HGA3 correspond to link2 and link3 respectively.

[0031] Read the external attitude data of the probe, obtain the transformation matrix from the high-gain antenna coordinate system of the probe to the satellite body coordinate system, and convert it into Euler angle format;

[0032] Multiply the obtained Euler angles with HGA2 and HGA3 respectively to obtain the correction values ​​corresponding to link3 and link2 in the satellite body coordinate system;

[0033] Add the correction values ​​corresponding to link3 and link2 to the first component of the centroid correction before projection to obtain the vector M of the centroid correction in the satellite body coordinate system before projection. SBF :

[0034] M SBF =HGA1+ANG2·HGA2+ANG3·HGA3

[0035] Among them, ANG2 and ANG3 are Euler angles;

[0036] Read the external attitude data of the probe, obtain the transformation matrix from the probe body coordinate system to the lunar center coordinate system, and compare the transformation matrix with the center of mass correction vector M in the satellite body coordinate system. SBF Multiply them together to obtain the center-of-mass correction vector M in the lunar coordinate system;

[0037] Satellite-based coordinates X s and the coordinates X of the ground tracking station t, subtract the two and divide by the modulus to obtain the unit vector of the ground tracking station pointing to the phase center of the high-gain antenna Multiply it with the center-of-mass correction vector M in the lunar coordinate system to obtain the center-of-mass correction value in the lunar coordinate system.

[0038] Optionally, the centroid correction value is expressed as:

[0039]

[0040] Among them, D r is the center of mass correction value, M diff represents the transformation matrix from the satellite body coordinate system to the moon center coordinate system, M SBF is the direction of the center of mass correction in the satellite body coordinate system, X s Represents the coordinates of the satellite, X t represents the coordinates of the ground tracking station, |X s -X t | represents the modulus of the subtraction of two vectors, Represents the unit vector pointing to the phase center of the high-gain antenna at the ground tracking station, m is the satellite mass, and M is the vector of the center of mass correction in the lunar coordinate system.

[0041] Optionally, the total acceleration of the detector also includes the acceleration caused by the atmospheric resistance on the detector, and the calculation formula is:

[0042]

[0043] Among them, a d represents the acceleration caused by the atmospheric drag on the probe, C D is the drag coefficient, ρ is the atmospheric density, is the detector speed, A i is the area of ​​sailboard i, θ vi is the angle between the normal vector of panel i and the probe velocity, N is the number of panels, and m is the mass of the satellite.

[0044] Optionally, the high-precision multi-sailboard model models the detector according to the material characteristics and external posture data of the detector itself, so as to construct a high-precision multi-sailboard model of the detector.

[0045] Optionally, the pressure generated by the solar radiation includes pressure generated by specular reflection and pressure caused by diffuse reflection.

[0046] Compared with the prior art, the present invention has the following beneficial effects:

[0047] The present invention calculates the centroid correction value by projecting the difference between the antenna offset vector and the centroid offset vector onto the unit vector pointing from the ground station to the tracking point, so as to further improve the orbit accuracy;

[0048] The present invention uses the external attitude data of the detector to establish a high-precision model of the detector, thereby calculating the light pressure and center of mass offset, and finally making corrections to obtain a more precise orbit with higher accuracy. This method establishes a high-precision model of the detector, replaces the simple force model in traditional orbit determination with the force calculated by the precise multi-sail board model, and performs continuous iterations to finally solve a high-precision precise orbit, providing basic support for the subsequent processing of other scientific instrument data. BRIEF DESCRIPTION OF THE DRAWINGS

[0049] Figure 1 This is a specific flow chart of the method for improving the orbit accuracy of a low-orbit detector of the present invention;

[0050] Figure 2 A schematic diagram of the invention for modeling a detector using a high-precision multi-sailboard model;

[0051] Figure 3 It is a schematic diagram of the high-gain antenna of the detector of the present invention and its mechanical arm;

[0052] Figure 4 It is a schematic diagram of the centroid correction from the ground tracking station to the high-gain antenna of the detector of the present invention. DETAILED DESCRIPTION

[0053] In order to make the technical means, creative features, objectives and effects achieved by the present invention easy to understand, the present invention is further explained below in conjunction with specific implementation methods.

[0054] The present invention aims to solve the problems of inaccurate light pressure and center of mass offset in the current precise orbit determination of lunar and Mars probes, and proposes a method for improving the orbit accuracy of low-orbit probes. Specifically, the external attitude data of the probe is used to establish a high-precision model of the probe, so as to calculate the light pressure and center of mass offset, and finally make corrections to obtain a more precise orbit with higher accuracy. The flowchart is shown in the attached figure. Figure 1 shown.

[0055] Embodiment 1 discloses a method for improving the orbit accuracy of a low-orbit detector, which specifically includes the following steps:

[0056] Step 1: Based on a high-precision multi-sailboard model, calculate the sum of the pressure generated by the incident sunlight and the pressure generated by solar radiation on each sailboard of the detector to obtain the acceleration generated by each sailboard;

[0057] Step 2, accumulating the acceleration generated by each sailboard to obtain the acceleration generated by the detector light pressure;

[0058] Step 3, based on the high-precision multi-sailboard model, calculate the acceleration of each sailboard of the probe caused by the planetary radiation pressure;

[0059] Step 4, accumulate the acceleration of each sailboard caused by the planetary radiation pressure to obtain the acceleration of the albedo and infrared radiation force on the detector;

[0060] Step 5, adding the acceleration generated by the light pressure of the detector and the acceleration of the albedo and infrared radiation force on the detector to obtain the total acceleration of the detector;

[0061] Step 6, perform numerical integration based on the total acceleration of the detector to obtain a calculated value;

[0062] Step 7, according to the external attitude data of the detector, calculate the centroid correction value from the phase center of the antenna to the centroid of the detector, add the centroid correction value to the calculated value, and obtain the corrected calculated value;

[0063] Step 8: Use the corrected calculated value to perform iterative calculations to obtain a high-precision precise orbit.

[0064] For step 1, the high-precision multi-sailboard model models the detector according to the material characteristics of the detector itself and the external attitude data, and constructs a high-precision multi-sailboard model of the detector. Taking the lunar probe as an example, since the lunar probe is composed of various materials, it is not possible to only consider the mirror reflection. Therefore, the present invention will C r It is split into two parts: specular reflection and diffuse reflection, and the satellite is modeled, using a variable area instead of the fixed satellite surface area A in the traditional model above to calculate a more precise light pressure.

[0065] The present invention uses 10 sailboards to model the lunar probe LRO, as shown in the attached figure. Figure 2 These include:

[0066] (1) Detector body SC (i.e. the rectangular block in the middle): +X, -X, +Y, -Y, +Z, -Z;

[0067] (2) Solar Array (front and back): +SA, -SA;

[0068] (3) High-gain antenna HGA (i.e. “pot”, front and back): +HGA, -HGA.

[0069] These three parts have their own coordinate systems, and each plate has a different reflection coefficient and normal vector, so as to calculate the light pressure of each plate.

[0070] For the multi-panel model of the lunar probe LRO, the present invention divides its light pressure into three parts: the pressure generated by the incident sunlight, the pressure generated by the specular reflection, and the pressure caused by the diffuse reflection. The sum of these three is the acceleration generated by the sunlight pressure on each sailboard.

[0071] In this embodiment, the acceleration calculation formula generated by each windsurfing board is:

[0072]

[0073] Among them, a i represents the acceleration generated by each sailboard, i is the number of the probe sailboard, there are N sailboards in total, C r is the solar radiation coefficient, P s is the solar pressure constant at 1 astronomical unit; A u is the length of an astronomical unit; r s is the distance from the sun to the satellite, A is the area of ​​the sailboard; θ is the angle between the normal vector of the sailboard and the direction of the sun; m is the mass of the satellite, δ is the diffuse reflectivity, and ρ is the specular reflectivity, all given as a percentage of the total incident radiation. is the normal vector of the sailboard, is the unit vector pointing from the satellite to the sun.

[0074] For step 2, the acceleration generated by each sailboard is accumulated to obtain the acceleration generated by the detector light pressure, which is calculated by the following formula:

[0075]

[0076] Among them, a s represents the acceleration generated by the light pressure of the detector, i is the number of the detector sailboard, there are N sailboards in total, C r is the solar radiation coefficient, P s is the solar pressure constant at 1 astronomical unit, A u is the length of an astronomical unit, r s is the distance from the sun to the satellite, θ i is the angle between the normal vector of sailboard i and the direction of the sun, A i is the area of ​​sailboard i, m is the mass of satellite, δ is diffuse reflectivity, ρ is specular reflectivity, is the normal vector of the sailboard i, is the unit vector pointing from the satellite to the sun.

[0077] For step 3, calculate the acceleration caused by the planetary radiation pressure: Calculate by the following formula: Based on the high-precision multi-sailboard model, the calculation formula for calculating the acceleration of each sailboard of the detector caused by the planetary radiation pressure is:

[0078]

[0079] in, represents the acceleration of each sailboard caused by the planetary radiation pressure, i is the number of the probe sailboard, there are N sailboards in total, j is the block divided by the planetary surface, there are J in total maxsurface blocks, G j is the radiation flux of the radiation source, A is the area of ​​the sailboard, m is the mass of the satellite, c is the speed of light, δ is the diffuse reflectivity, ρ is the mirror reflectivity, θ j represents the angle between the normal vector of the sailboard and the vector of the block j pointing to the planet surface, is the normal vector of the sailboard, is the unit vector pointing from the satellite to the source of the radiation.

[0080] For step 4, the acceleration of each sailboard caused by the planetary radiation pressure is accumulated to obtain the acceleration calculation formula of the albedo and infrared radiation force on the detector, which is calculated by the following formula:

[0081]

[0082] Among them, a a It represents the acceleration of the albedo and infrared radiation force on the probe, i is the number of the probe sailboard, there are N sailboards in total, j is the block divided by the planetary surface, there are J in total max surface blocks, G j is the radiation flux of the radiation source, A i is the area of ​​sailboard i, m is the mass of the satellite, c is the speed of light, δ is the diffuse reflectivity, ρ is the mirror reflectivity, and ρ i is the mirror reflectivity of sailboard i, θ ij represents the angle between the normal vector of sailboard i and the vector of sailboard pointing to block j on the surface of Mars. If θ ij Greater than 90 。 , then the acceleration is 0, is the normal vector of the sailboard, is the unit vector pointing from the satellite to the source of the radiation.

[0083] In this embodiment, in step 6, numerical integration is performed according to the total acceleration of the detector to obtain a calculated value, which is common knowledge in the art and will not be described in detail here.

[0084] For step 7, based on the high-precision multi-sail model and the external attitude data of the detector, the center of mass correction value from the phase center of the antenna to the center of mass of the detector is calculated, including:

[0085] The vector from the detector's center of mass to the high-gain antenna's phase center is calculated using a high-precision multi-sail model;

[0086] Perform coordinate frame transformation on vectors;

[0087] Project the converted vector to the direction of the ground tracking station pointing to the phase center of the high-gain antenna, and calculate the dot product of the two vectors to obtain the centroid correction value. This is achieved by the following steps:

[0088] Taking the lunar probe LRO as an example, due to its special shape and long antenna, the distance between these two points can often reach several meters, which has a certain impact on the final orbit determination result. Therefore, centroid correction is required to calculate the correction from the phase center of the antenna to the centroid of the probe.

[0089] Since the high-gain antenna is controlled and continuously rotated by multiple robotic arms, and the lunar probe LRO has a special structure with multiple robotic arms controlling the multi-degree-of-freedom rotation of the antenna, this correction cannot be made by simply adding a fixed constant.

[0090] Using the multi-sail model established in the first problem, the vector from the satellite centroid to the phase center of the high-gain antenna can be calculated. After multiple coordinate frame transformations, this vector is finally projected to the direction of the ground tracking station (measurement station) pointing to the phase center of the high-gain antenna. The dot product of the two vectors is calculated to obtain the centroid correction value. The schematic diagram is attached. Figure 4 shown.

[0091] The centroid correction must be performed every time the distance (light travel time) is calculated, so we first need to obtain the current ephemeris time et and convert it into the number of seconds after the J2000 epoch (seconds past J2000) for easy calculation. Then, based on the multi-sail model established earlier and the basic parameters of the lunar probe LRO, we can obtain its initial offset vector and correspond each component of the offset vector to the actual multiple high-gain antenna manipulators of the lunar probe LRO, that is:

[0092]

[0093] Among them, offset is the initial offset vector, HGA1 corresponds to link1, and since link1 is connected to the detector body, it is directly regarded as the first component of the centroid correction before projection (projection has not been performed yet).

[0094] Then, since this offset vector is in the LRO high-gain antenna coordinate system, we need to convert it to the LRO body coordinate system (SBF) for subsequent calculations. By reading the LRO external attitude data, we can obtain the transformation matrix from the high-gain antenna coordinate system to the body coordinate system and the corresponding Euler angles.

[0095] After calculating the Euler angle from the high-gain antenna coordinate system to the satellite body coordinate system, it can be multiplied by the offset vector HGA3 corresponding to link3 to obtain the offset vector corresponding to link3 in the satellite body coordinate system. Similarly, multiply the Euler angle from the high-gain antenna coordinate system to the satellite body coordinate system by the offset vector HGA2 corresponding to link2 to obtain the offset vector corresponding to link2 in the satellite body coordinate system. Add the two offset vectors obtained to the offset vector corresponding to link1 to obtain the center of mass correction vector M in the satellite body coordinate system. SBF (Not projected):

[0096] M SBF =HGA1+ANG2·HGA2+ANG3·HGA3

[0097] Since the precise orbit determination finally obtains the coordinates in the J2000 coordinate system at the center of the moon, it is necessary to transform this vector into a coordinate system again. By reading the external attitude data of the lunar probe LRO, the transformation matrix from the satellite body coordinate system to the J2000 coordinate system can be calculated and compared with the vector M of the center of mass correction in the satellite body coordinate system. SBF Multiply them together to get the center of mass correction vector M in the J2000 coordinate system. J2000 .

[0098] Finally, this vector needs to be projected to the direction of the ground tracking station (measurement station) pointing to the phase center of the high-gain antenna. When calculating the speed measurement, distance measurement and other calculation values, the coordinates of the satellite and the ground tracking station will be calculated. Subtracting and normalizing the two will give the unit vector of the ground tracking station pointing to the phase center of the high-gain antenna. Finally, it is combined with the center of mass correction vector M in the J2000 coordinate system. J2000 Multiply them together to get the center of mass correction value D in the J2000 coordinate system. r .

[0099] In this embodiment, the centroid correction value is expressed as:

[0100]

[0101] Among them, D r is the center of mass correction value, M diff represents the transformation matrix from the satellite body coordinate system to the moon center coordinate system, M SBF is the vector of the center of mass correction in the satellite body coordinate system, X s Represents the coordinates of the satellite, X t represents the coordinates of the ground tracking station, |X s -X t | represents the modulus of the subtraction of two vectors, Represents the unit vector pointing to the phase center of the high-gain antenna at the ground tracking station, m is the satellite mass, and M is the vector of the center of mass correction in the lunar coordinate system.

[0102] Example 2

[0103] The difference from Example 1 is that when the probe is a Mars probe, a multi-sailboard model also needs to be established for the Mars probe. The acceleration formula generated by the light pressure perturbation force it receives is basically the same as that of the lunar probe.

[0104] In addition to the light pressure generated by direct sunlight shining on the detector, the detector is also affected by the albedo and other factors. The acceleration caused by the albedo is relatively complex and difficult to calculate directly, so an approximate method is used: with the satellite's subsatellite point as the center, the planetary surface seen from the satellite can be divided into several rings. The central part is called the central crown, and the rings outside the central crown are further divided into concentric ring segments. The albedo and infrared radiation pressure of the central crown and each segment on the satellite are considered to be concentrated at the center point of the block. The radiation pressure of the central crown has only a radial component on the satellite, and the radiation pressure of each small segment on the satellite has radial, lateral and normal components.

[0105] Taking the Mars rover as an example, for each segment and each sail, calculate the acceleration due to the planetary radiation pressure:

[0106]

[0107] Then sum them up to get the Mars albedo and infrared radiation force on MGS:

[0108]

[0109] Where G is the radiation flux of the radiation source (W / m 2 ), A is the surface area of ​​the sailboard, m is the mass of the satellite, c is the speed of light, δ is the diffuse reflectivity, ρ is the mirror reflectivity, is the panel normal vector, is the unit vector pointing to the radiation source. i is the number of the probe sailboard, there are N sailboards in total, and J is the block divided on the surface of Mars, there are J in total. max Surface patches. ij角 degrees represents the angle between the normal vector of the sailboard and the block vector pointing to the surface of Mars. If θ ij If it is greater than 90°, the acceleration is 0, because the sailboard i cannot observe the corresponding surface block j.

[0110] Because there is an atmosphere on Mars, unlike the lunar probe LRO, the Mars probe MGS also needs to consider the atmospheric drag on the probe, which can be calculated using the following formula:

[0111]

[0112] Among them, a d represents the acceleration caused by the atmospheric drag on the probe, C D is the drag coefficient, ρ is the atmospheric density, is the detector speed, A i is the area of ​​sailboard i, θ vi is the angle between the normal vector of panel i and the probe velocity, N is the number of panels, and m is the mass of the satellite.

[0113] Finally, these forces calculated by the precise multi-sail model are used to replace the simple force model in traditional orbit determination, and continuous iterations are performed to ultimately solve a high-precision orbit.

[0114] The method described in the present invention can calculate the precise light pressure and center of mass offset received by the detector, and finally make corrections to obtain a more precise orbit with higher accuracy, thus providing basic support for the subsequent processing of data from other scientific instruments.

[0115] The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the technical principles of the present invention. These improvements and modifications should also be regarded as the scope of protection of the present invention.

Claims

1. A method for improving the orbit accuracy of a low-orbit detector, characterized in that: The method comprises: Based on a high-precision multi-board model, the sum of the pressure generated by the incident sunlight and the pressure generated by solar radiation on each board of the detector is calculated to obtain the acceleration generated by each board. Accumulate the acceleration generated by each sailboard to obtain the acceleration generated by the detector light pressure; Based on a high-precision multi-panel model, the acceleration of each panel of the probe caused by the planetary radiation pressure is calculated; The acceleration of each sailboard caused by the planetary radiation pressure is accumulated to obtain the acceleration of the albedo and infrared radiation force on the probe; The total acceleration of the detector is obtained by adding the acceleration generated by the light pressure of the detector and the acceleration of the albedo and infrared radiation force on the detector. According to the total acceleration of the detector, numerical integration is performed to obtain the calculated value; According to the external attitude data of the detector, the centroid correction value from the phase center of the antenna to the centroid of the detector is calculated, and the centroid correction value is added to the calculated value to obtain the corrected calculated value; The corrected calculated values ​​are used for continuous iterative calculations to obtain a high-precision precise orbit.

2. The method for improving the orbital accuracy of a low-orbit detector according to claim 1, characterized in that: The acceleration generated by each windsurfing board is calculated as: Among them, a i represents the acceleration generated by each sailboard, i is the number of the probe sailboard, there are N sailboards in total, C r is the solar radiation coefficient, P s is the solar pressure constant at 1 astronomical unit, A u is the length of an astronomical unit; r s is the distance from the sun to the satellite; A is the area of ​​the sailboard; θ is the angle between the normal vector of the sailboard and the direction of the sun; m is the mass of the satellite, δ is the diffuse reflectivity, ρ is the specular reflectivity, is the normal vector of the sailboard; is the unit vector pointing from the satellite to the sun.

3. The method for improving the orbital accuracy of a low-orbit detector according to claim 1, characterized in that: The acceleration generated by each sailboard is accumulated to obtain the calculation formula of the acceleration generated by the detector light pressure: Among them, a s represents the acceleration generated by the light pressure of the detector, i is the number of the detector sailboard, there are N sailboards in total, C r is the solar radiation coefficient, P s is the solar pressure constant at 1 astronomical unit, A u is the length of the astronomical unit, r s is the distance from the sun to the satellite, θ i is the angle between the normal vector of sailboard i and the direction of the sun, A i is the area of ​​sailboard i, m is the mass of satellite, δ is diffuse reflectivity, ρ is specular reflectivity, is the normal vector of the sailboard i, is the unit vector pointing from the satellite to the sun.

4. The method for improving the orbital accuracy of a low-orbit detector according to claim 1, characterized in that: The formula for calculating the acceleration of each sailboard of the detector caused by the planetary radiation pressure based on the high-precision multi-sailboard model is: in, represents the acceleration of each sailboard caused by the planetary radiation pressure, i is the number of the probe sailboard, there are N sailboards in total, j is the block divided by the planetary surface, there are J in total max surface blocks, G j is the radiation flux of the radiation source, A is the area of ​​the sailboard, m is the mass of the satellite, c is the speed of light, δ is the diffuse reflectivity, ρ is the mirror reflectivity, θ j represents the angle between the normal vector of the sailboard and the vector of the block j pointing to the planet surface, is the normal vector of the sailboard, is the unit vector pointing from the satellite to the source of the radiation.

5. The method for improving the orbital accuracy of a low-orbit detector according to claim 1, characterized in that: The acceleration of each sailboard caused by the planetary radiation pressure is accumulated to obtain the acceleration calculation formula of the albedo and infrared radiation force on the detector: Among them, a a It represents the acceleration of the albedo and infrared radiation force on the probe, i is the number of the probe sailboard, there are N sailboards in total, j is the block divided by the planetary surface, there are J in total max surface blocks, G j is the radiation flux of the radiation source, A i is the area of ​​sailboard i, m is the mass of the satellite, c is the speed of light, δ is the diffuse reflectivity, ρ is the mirror reflectivity, and ρ i is the mirror reflectivity of sailboard i, θ ij represents the angle between the normal vector of sailboard i and the vector of sailboard pointing to block j on the surface of Mars. If θ ij Greater than 90 。 , then the acceleration is 0, is the normal vector of the sailboard, is the unit vector pointing from the satellite to the source of the radiation.

6. The method for improving the orbital accuracy of a low-orbit detector according to claim 1, characterized in that: The method of calculating the center of mass correction value from the phase center of the antenna to the center of mass of the detector based on the high-precision multi-sailboard model and the external attitude data of the detector includes: Based on the detector design data, the initial offset vector of the detector is obtained: Wherein, offset is the initial offset vector, HGA1 corresponds to link1. Since link1 is connected to the detector body, it is directly regarded as the first component of the centroid correction before projection. HGA2 and HGA3 correspond to link2 and link3 respectively. Read the external attitude data of the probe, obtain the transformation matrix from the high-gain antenna coordinate system of the probe to the satellite body coordinate system, and convert it into Euler angle format; Multiply the obtained Euler angles with HGA2 and HGA3 respectively to obtain the correction values ​​corresponding to link3 and link2 in the satellite body coordinate system; Add the correction values ​​corresponding to link3 and link2 to the first component of the centroid correction before projection to obtain the vector M of the centroid correction in the satellite body coordinate system before projection. SBF : M SBF =HGA1+ANG2·HGA2+ANG3·HGA3 Among them, ANG2 and ANG3 are Euler angles; Read the external attitude data of the probe, obtain the transformation matrix from the probe body coordinate system to the lunar center coordinate system, and compare the transformation matrix with the vector M of the center of mass calibration in the satellite body coordinate system. SBF Multiply them together to get the center-of-mass calibration vector M in the lunar coordinate system: Satellite-based coordinates X s and the coordinates X of the ground tracking station t , subtract the two and divide by the modulus to obtain the unit vector of the ground tracking station pointing to the phase center of the high-gain antenna Multiply it with the center-of-mass correction vector M in the lunar coordinate system to obtain the center-of-mass correction value in the lunar coordinate system.

7. The method for improving the orbital accuracy of a low-orbit detector according to claim 6, characterized in that: The centroid correction value is expressed as: Among them, D r is the center of mass correction value, M diff represents the transformation matrix from the satellite body coordinate system to the moon center coordinate system, M SBF is the direction of the center of mass correction in the satellite body coordinate system, X s Represents the coordinates of the satellite, X t represents the coordinates of the ground tracking station, |X s -X t | represents the module of two directions subtracted from each other, Represents the unit vector pointing to the phase center of the high-gain antenna at the ground tracking station, m is the satellite mass, and M is the vector of the center of mass correction in the lunar coordinate system.

8. The method for improving the orbital accuracy of a low-orbit detector according to claim 1, characterized in that: The total acceleration of the detector also includes the acceleration caused by the atmospheric resistance of the detector, and the calculation formula is: Among them, a d represents the acceleration caused by the atmospheric drag on the probe, C D is the drag coefficient, ρ is the atmospheric density, is the detector speed, A i is the area of ​​sailboard i, θ vi is the angle between the normal vector of panel i and the probe velocity, N is the number of panels, and m is the mass of the satellite.

9. The method for improving the orbital accuracy of a low-orbit detector according to claim 1, characterized in that: The high-precision multi-sailboard model models the detector according to the material characteristics of the detector itself and external posture data, and constructs a high-precision multi-sailboard model of the detector.

10. The method for improving the orbital accuracy of a low-orbit detector according to claim 1, characterized in that: The pressure generated by the solar radiation includes the pressure generated by specular reflection and the pressure caused by diffuse reflection.

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