Static orbit microwave sounding satellite payload-star sensitive thermal deformation on-orbit decoupling calibration device

Abstract

The application provides a static orbit microwave detection satellite load-star sensitive thermal deformation on-orbit decoupling calibration device, the microwave load includes a main reflector, a first secondary reflector, a second secondary reflector and a quasi-optical system;The star sensor is installed on the side plate of the microwave detection satellite platform;The laser scanning measurement camera monitors the on-orbit thermal deformation law of each reflector itself by measuring the coplanar laser target space position on the main reflector, the first secondary reflector and the second secondary reflector;The common point conversion target is used for realizing the benchmark transmission of the laser measurement results of the coplanar laser targets of each reflector;The coplanar laser target is used for establishing the reflector body coordinate system, and the on-orbit calibration result and the ground calibration result are directly compared to obtain the on-orbit thermal deformation law of the reflector itself;The infrared camera is installed with the quasi-optical frame as a common benchmark, and the overall thermal deformation law of the microwave load-star sensor at different landmark observation moments is monitored through the on-orbit infrared landmark observation result.

Description

Technical Field

[0001] This invention relates to the field of satellite microwave remote sensing technology, specifically to an on-orbit decoupling calibration device for payload-star-sensor thermal deformation of geostationary orbit microwave sounding satellites, and more specifically to an on-orbit decoupling calibration system for payload-star-sensor thermal deformation of geostationary orbit microwave sounding satellites for real-time on-orbit imaging, navigation, and registration. Background Technology

[0002] Geostationary microwave sounding satellites combine the high timeliness of geostationary satellites with the ability of microwaves to penetrate clouds and rain atmospheres, which can greatly enhance the monitoring and early warning capabilities for sudden disasters such as typhoons and rainstorms.

[0003] When geostationary microwave sounding satellites operate in orbit, the complex external heat flow environment causes structural thermal deformation of the large reflector antenna and star sensor on the microwave payload due to drastic temperature changes. This leads to the antenna's electrical axis deviating from its nominal orientation, causing the remote sensing data to drift relative to the ideal grid position, severely affecting the accuracy and timeliness of subsequent quantitative applications. Geostationary optical remote sensing satellites can compensate for the optical axis pointing deviation caused by the thermal deformation of the scanning mirror and star sensor through a landmark observation-based optical axis pointing correction model, achieving real-time navigation and registration of optical remote sensing data and obtaining high-precision Earth positioning results. Compared to geostationary optical remote sensing satellites, geostationary microwave sounding satellites have a limited observation area and slower scanning speed, making it difficult to ensure timely landmark observation experiments. Therefore, how to achieve decoupling calibration of "payload-star sensor" thermal deformation on-board to ensure the stability and continuity of imaging navigation and registration operations when landmark observation data is missing remains a key challenge.

[0004] Based on a search of existing technologies:

[0005] Patent document CN110345923A (application number: 201810307554.4) discloses an antenna primary and secondary reflector pose measurement system and a secondary reflector pose adjustment method. This method uses an industrial camera fixed to the primary reflector to continuously photograph the secondary reflector, measuring the relative pose relationship between the secondary reflector and the camera in real time. The relative pose relationship between the primary and secondary reflectors is then calculated based on the known relative pose relationship between the primary and secondary reflectors. However, the external thermal environment of geostationary orbit differs from the ground environment, and the relative pose relationship between the primary and secondary reflectors undergoes periodic changes, making it impossible to accurately calibrate the changes in the relative pose relationship between the primary and secondary reflectors. Furthermore, this method cannot calibrate the on-orbit thermal deformation of the star sensor, and therefore is not suitable for the on-orbit decoupling calibration of the thermal deformation of the "payload-star sensor" on the geostationary microwave detection satellite.

[0006] Patent document CN108879106A (application number: 201810620306.5) discloses a method for real-time adjustment of the surface shape of a dual-reflector antenna based on best-fit. This method uses a laser measurement device installed on the back of the sub-reflector to measure the surface shape of the main reflector. While detecting the surface shape of the main reflector, the pose matching of the sub-reflector is achieved according to the best-fit algorithm. The drawback of this method is that the laser measurement system is attached to the sub-reflector, and the laser probe needs to be moved by a ring guide rail installed on the back of the sub-reflector to achieve real-time measurement of the surface shape of the main reflector. Therefore, it is mainly suitable for closed-loop measurement and control of dual-reflector antennas, and is not suitable for antenna systems with more reflectors and more complex configurations. Furthermore, this method cannot calibrate the on-orbit thermal deformation of the star sensor, making it difficult to apply to the on-orbit decoupling calibration of thermal deformation of the "payload-star sensor" of the geostationary microwave sounding satellite.

[0007] Patent document CN110553667A (application number: 201910857315.0) discloses a method for compensating for the thermal deformation of a star sensor. This method obtains a temperature-thermal deformation fitting formula for the star sensor through ground-based thermal deformation experiments. In actual on-orbit operation, it acquires temperature telemetry data from the star sensor's temperature measurement points and, based on the temperature-thermal deformation fitting formula, obtains the attitude transformation matrix from the actual installation coordinate system to the theoretical installation coordinate system of the star sensor. However, since ground-based thermal deformation experiments cannot accurately simulate the complex thermal environment of a geostationary orbit, the practical operability of this method for on-orbit calibration of the thermal deformation of star sensors on the aforementioned geostationary microwave detection satellite lacks persuasiveness.

[0008] Patent document CN111380567A (application number: 202010214353.7) discloses an on-orbit thermal deformation correction method for satellite star sensor pointing. This method obtains the relationship between the temperature of a large planar antenna load and the star sensor pointing through ground measurements, and corrects the star sensor pointing in real time in orbit by measuring the antenna temperature. Since the large planar antenna load has a short on-orbit operating time, the antenna temperature change is mainly caused by power-on. The star sensor pointing correction model obtained from ground-based power-on tests can be used in the on-orbit operating state. However, for geostationary microwave detection satellites, the thermal deformation state of the star sensor during ground thermal tests differs significantly from the thermal deformation state during on-orbit operation. Therefore, this method cannot meet the on-orbit calibration requirements for star sensor thermal deformation.

[0009] Patent document CN112162158A (application number: 202010913585.1) discloses a method and apparatus for evaluating antenna pointing mismatch of an on-orbit terahertz detector. It uses cosmic radio sources as a space reference standard to establish a model of the influence of antenna thermal deformation on the electrical axis pointing deviation of the terahertz detector. By observing a sufficient number of cosmic radio sources in a short period, the antenna electrical axis pointing deviation is solved. This method belongs to the overall thermal deformation calibration method of "microwave payload-star sensor," and cannot achieve on-orbit decoupling calibration of the thermal deformation of the "payload-star sensor" of the geostationary microwave detection satellite. Summary of the Invention

[0010] To address the shortcomings of existing technologies, the purpose of this invention is to provide an on-orbit decoupling calibration device for geostationary microwave detection satellite payload-star-sensitive thermal deformation.

[0011] According to the present invention, a geostationary microwave detection satellite payload-star sensor thermal deformation on-orbit decoupling calibration device includes: a microwave payload, a star sensor, a laser scanning measurement camera, a common point conversion target, a coplanar laser target, and an infrared camera;

[0012] The microwave payload includes: a primary antenna reflector, a first secondary reflector, a second secondary reflector, an antenna strut, and a quasi-optical system; the primary antenna reflector, the first secondary reflector, and the second secondary reflector are mounted on the top plate of the satellite platform via the antenna strut; the quasi-optical system is mounted and embedded in the satellite platform via a quasi-optical frame;

[0013] The microwave payload and the satellite platform are installed at an angle according to preset requirements;

[0014] The star sensor is installed on the side plate of the geostationary microwave detection satellite platform to ensure that the satellite pointing reference is consistent with the attitude control reference.

[0015] The laser scanning measurement cameras are respectively mounted on a pair of anti-bracing rods and below the collimation frame;

[0016] The infrared camera and the quasi-optical frame are mounted together with a common reference.

[0017] The common point conversion target is affixed to the central area of ​​a pair of anti-bracing rods and the main anti-shape surface;

[0018] The coplanar laser targets are arranged on three reflective surfaces.

[0019] Preferably, the laser scanning measurement camera includes two sets. The laser scanning measurement camera installed on the first set of countersupport rods monitors the on-orbit thermal deformation of the main countersupport by measuring the spatial position of the target points on the main countersupport. The laser scanning measurement camera installed below the quasi-optical frame monitors the on-orbit thermal deformation of the first and second countersupports by measuring the spatial position of the target points on the first and second countersupports.

[0020] Preferably, the infrared camera selects a working band that meets preset requirements to acquire infrared landmark images of cloudless or less cloudy areas. Based on the infrared landmark observation results, the three-axis pointing deviation angle of the infrared camera's optical axis relative to the star sensor support coordinate system is calculated, and then the overall thermal deformation law of the microwave payload-star sensor at different landmark observation times is calibrated on orbit.

[0021] Preferably, the common point conversion target is arranged in the central area of ​​a pair of anti-support rods and the main anti-reflection surface, respectively. The common point conversion target is used to realize the reference transfer of the laser measurement results of the coplanar laser target of each reflecting surface, and to unify the measurement data of the two sets of laser scanning measurement cameras under the main anti-reflection design coordinate system.

[0022] Preferably, the coplanar laser targets are arranged on three reflective surfaces to establish a body coordinate system for each reflective surface. By directly comparing the on-orbit calibration results of the body coordinate system with the ground calibration results, the on-orbit thermal deformation of each reflective surface can be obtained.

[0023] Preferably, the infrared camera adopts an antenna-converging lens-rear optical path assembly. The converging lens is mounted on a support rod near the second secondary mirror, and the rear optical path is mounted on the collimating frame and the connecting rod of the second secondary mirror, sharing a common reference with the collimating frame.

[0024] Preferably, the main reflector of the microwave load is a 5m diameter parabolic rotating surface, the secondary reflector is a 1m diameter hyperboloid rotating surface, and the third secondary reflector is a 1.4m diameter planar surface.

[0025] Preferably, after the microwave load is installed at an angle, the main inverse design coordinate system is positioned relative to the Earth's line of sight +z. M With the satellite body coordinate system +z B The axis forms an angle of 23.997°.

[0026] Preferably, the star sensor is mounted on a star sensor bracket, and the star sensor bracket and the satellite platform are integrated into a mechanical and thermal structure.

[0027] Preferably, the distance-angle spherical coordinates measured by the laser scanning measurement camera in the laser measurement coordinate system are converted into Cartesian three-dimensional spatial coordinates, and the on-orbit thermal deformation is monitored based on the Cartesian three-dimensional spatial coordinates.

[0028] Compared with the prior art, the present invention has the following beneficial effects:

[0029] 1. The geostationary microwave sounding satellite payload-star sensor thermal deformation on-orbit decoupling calibration device provided by this invention addresses the problems of limited observation area, slow scanning transmission speed, and inability to ensure timely landmark observation experiments of geostationary microwave sounding satellites. It designs an error decoupling optical path in conjunction with a laser scanning measurement camera and an infrared camera. The laser scanning measurement camera obtains the on-orbit thermal deformation law of each reflector of the microwave payload, and the infrared camera obtains the overall coupled thermal deformation law of the "microwave payload-star sensor". After removing the thermal deformation law of each reflector of the microwave payload from the overall thermal deformation law of the "microwave payload-star sensor", the on-orbit thermal deformation law of the star sensor can be obtained. This provides a new approach to ensure the stability and continuity of imaging navigation and registration services when special observation needs cause the lack of landmark observation data during the on-orbit operation of geostationary microwave sounding satellites.

[0030] 2. This invention proposes an on-orbit decoupling calibration device for geostationary microwave sounding satellite payload-star-sensor thermal deformation, which ensures the stability and continuity of imaging navigation and registration services during the on-orbit operation of geostationary microwave sounding satellites when special observation requirements cause the loss of infrared landmark observation data.

[0031] 3. The synchronous monitoring results of the microwave payload-star sensor overall thermal deformation law and the thermal deformation law of each reflective surface of the microwave payload can accurately calibrate the on-orbit thermal deformation law of the star sensor, thus solving the problem of on-orbit decoupling calibration of thermal deformation of geostationary microwave sounding satellite payload-star sensor. It meets the requirements of all-weather, space-ground integrated high-precision imaging, navigation and registration for geostationary microwave sounding satellites. At the same time, the invention also has the advantages of easy engineering implementation and strong operability, and has a certain degree of versatility. It can be widely used in various geostationary microwave sounding satellite on-orbit deformation measurement and control subsystems. Attached Figure Description

[0032] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0033] Figure 1 This is a schematic diagram of an on-orbit decoupling calibration device for a geostationary microwave probe satellite payload-star-sensitive thermal deformation.

[0034] Figure 2 This is a schematic diagram of the infrared positioning camera antenna-converging lens-rear optical path assembly.

[0035] Figure 3 This is a flowchart for the transfer of measurement references.

[0036] Figure 4 This is a schematic diagram of the body coordinate system of the antenna reflector.

[0037] Figure 5 Flowchart of the on-orbit decoupling calibration method for geostationary microwave probe satellite payload-star-sensitive thermal deformation.

[0038] 1-Main reflector; 2-First secondary reflector; 3-Second secondary reflector; 4-Infrared camera; 5-First laser scanner; 6-Second laser scanner; 7-Optical alignment frame; 8-Second secondary counter-support rod; 9-Second secondary counter-support rod; 10-Rear optical path assembly; 11-Converging lens; 12-Light shield. Detailed Implementation

[0039] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0040] Example 1

[0041] Before proceeding with the specific embodiments, the coordinate systems involved are explained as follows:

[0042] Satellite body coordinate system: Origin of coordinate system B For the satellite's center of mass, +x B The axis points in the direction of satellite flight, +z B The axis points to the sub-star point, +y B axis and +x B +z B The axes form a right-handed rectangular coordinate system;

[0043] Main inverse design coordinate system: Origin of the coordinate system M For the geometric vertex of the parabolic surface of revolution, +z M The axis coincides with the focal axis of the paraboloid of revolution and points in the positive normal direction to the geometric vertex of the paraboloid of revolution, +x M Axis perpendicular + z M The axis points to a characteristic direction of the parabolic surface of revolution (with respect to the satellite's body coordinate system +z). B (The axes form an angle of 66.003°), +y M axis and +z M +x M The axes form a right-handed rectangular coordinate system;

[0044] Infrared camera coordinate system: Origin of the coordinate system C Located at the center of the camera's focal plane, +z C The axis represents the direction of the optical axis, +x C axis and +y C The axes are two orthogonal references defined on the camera's focal plane;

[0045] Star-type support coordinate system: Origin of the coordinate system O s Located at the center of the star-sensitive photosensitive surface, +z S The axis points in the direction of the star sensor's principal optical axis, +x S axis and +y S The axes are two orthogonal references defined on the star-sensitive photosensitive surface.

[0046] like Figures 1 to 2 As shown, the present invention provides an on-orbit decoupling calibration device for thermal deformation of a geostationary microwave sounding satellite payload-star sensor. This device achieves real-time monitoring of the on-orbit thermal deformation of the three-reflector solid-plane antenna by installing a laser scanning measurement camera on the microwave payload. Simultaneously, it performs high-precision calibration of the overall on-orbit thermal deformation of the "microwave payload-star sensor" using an infrared camera. Based on the synchronous monitoring results of the overall thermal deformation law of the "microwave payload-star sensor" and the thermal deformation law of the microwave payload itself, the on-orbit thermal deformation law of the star sensor is calculated. This achieves the decoupling calibration of the thermal deformation of the geostationary microwave sounding satellite payload-star sensor, meeting its requirements for all-weather, integrated satellite-ground high-precision imaging, navigation, and registration.

[0047] The geostationary orbit microwave detection satellite payload-star sensor thermal deformation on-orbit decoupling calibration device includes: microwave payload, star sensor, laser scanning measurement camera, common point conversion target, coplanar laser target and infrared camera;

[0048] The microwave payload includes an antenna main reflector 1 (referred to as the main reflector), a first secondary reflector 2 (referred to as the secondary reflector), a second secondary reflector 3 (referred to as the secondary reflector), an antenna strut, and a quasi-optical system. The three reflectors are mounted on the top plate of the satellite platform via the antenna strut. The quasi-optical system is mounted and embedded in the satellite platform via the quasi-optical frame 7. The microwave payload and the satellite platform are installed at an angle. The star sensor is mounted on the side plate of the geostationary microwave detection satellite platform, and is aligned with the satellite reference as much as possible to ensure that the satellite pointing reference is consistent with the attitude control reference. The laser scanning measurement camera consists of two sets, including a first laser scanner 5 and a second laser scanner 6, which are respectively positioned below the secondary reflector strut 8 and the quasi-optical frame 7. The common point conversion target is pasted in the central area of ​​the secondary reflector strut 8 and the main reflector 1. The coplanar laser targets are respectively arranged on the three reflectors. The infrared camera 4 is mounted with the quasi-optical frame 7 on the same reference. Based on the error decoupling optical path, the on-orbit thermal deformation law of each reflective surface of the microwave load is obtained by the laser scanning measurement camera, and the overall coupled thermal deformation law of "microwave load-star sensor" is obtained by the infrared camera 4. After removing the thermal deformation law of each reflective surface of the microwave load from the overall thermal deformation law of "microwave load-star sensor", the on-orbit thermal deformation law of the star sensor itself can be obtained.

[0049] The laser scanning measurement camera includes: a first laser scanner 5 and a second laser scanner 6; the first laser scanner 5 is mounted on a set of countersuppression rods 8, and the first laser scanner 5 mounted on the set of countersuppression rods 8 monitors the on-orbit thermal deformation of the main reflector by measuring the spatial position of the target points on the main reflector; the second laser scanner 6 is mounted below the quasi-optical frame 7, and the second laser scanner 6 mounted below the quasi-optical frame 7 monitors the on-orbit thermal deformation of the first and second sub-reflectors by measuring the spatial position of the target points on the first and second sub-reflectors;

[0050] The common point conversion targets are respectively arranged in the central area of ​​the 8 (6) reverse struts and the 1-shaped surface of the main reflection (4). The common point conversion targets are used to realize the reference transfer of the laser measurement results of the coplanar laser target of each reflecting surface, and to unify the measurement data of the two sets of laser scanning measurement cameras under the design coordinate system of the main reflection.

[0051] The coplanar laser targets are arranged on three reflective surfaces (four targets are arranged on each reflective surface) to establish the body coordinate system of each reflective surface. By directly comparing the on-orbit calibration results of the body coordinate system with the ground calibration results, the on-orbit thermal deformation of each reflective surface can be obtained.

[0052] The infrared camera 4 adopts an "antenna-converging lens-rear optical path" assembly. The converging lens 11 is mounted on the support rod 9 near the second secondary mirror. The rear optical path assembly 10 is mounted on the collimating frame 7 and the second secondary mirror support rod 9, sharing the same reference with the collimating frame 7. The light shield 12 is mounted on the converging lens 11.

[0053] Specifically, the main reflector 1 of the stationary orbit microwave load is a 5m diameter parabolic rotating surface, the secondary reflector 2 is a 1m diameter hyperboloid rotating surface, and the secondary reflector 3 is a 1.4m diameter plane.

[0054] Specifically, after the microwave load is installed at an angle, the main inverse design coordinate system's line of sight to the ground +z... M With the satellite body coordinate system +z B The axis forms an angle of 23.997°.

[0055] Specifically, the star sensor is mounted on a star sensor bracket, and the star sensor bracket and the satellite platform are integrated into a mechanical and thermal structure.

[0056] Specifically, the laser scanning measurement camera consists of a two-dimensional scanning mechanism, a laser optical path, a ranging optical path, a detector, a light shield, etc., with a wavelength of 1500 / 1560nm. The distance-angle spherical coordinates measured in the laser measurement coordinate system can be converted into Cartesian three-dimensional spatial coordinates.

[0057] Specifically, the infrared camera 4 uses the 10μm working band with strong atmospheric penetration to acquire 4km infrared landmark images of cloudless or low-cloud areas. Based on the infrared landmark observation results, the three-axis pointing deviation angle of the infrared camera's optical axis relative to the star sensor support coordinate system is calculated, and then the overall thermal deformation law of the "microwave payload-star sensor" at different landmark observation times is calibrated on orbit.

[0058] Example 2

[0059] Example 2 is a preferred example of Example 1.

[0060] This invention provides an on-orbit decoupling calibration method for geostationary microwave sounding satellite payload-star-sensor thermal deformation, which employs the aforementioned on-orbit decoupling calibration device for geostationary microwave sounding satellite payload-star-sensor thermal deformation to perform the following steps:

[0061] Step 1: Establish a body coordinate system for the reflecting surface using a coplanar laser target, and calibrate the relationship between each body coordinate system and the corresponding laser measurement coordinate system on the ground;

[0062] Specifically, such as Figure 3 As shown, the body coordinate system is established using:

[0063] A corresponding body coordinate system is generated using four two-dimensional coplanar target points on the reflector surface, allowing the on-orbit thermal deformation of the reflector surface to be reflected by changes in the position and attitude (or pose) of the body coordinate system. Let A, B, C, and D be the four two-dimensional coplanar targets on the reflector surface. The spatial coordinates of the four coplanar targets in the laser measurement coordinate system, measured by the laser scanning measurement device, are A(x, y, z), B(x, y, z), C(x, y, z), and D(x, y, z), respectively. The center Q of the four coplanar targets is taken as the origin of the body coordinate system of the reflector surface, as shown in the following equation:

[0064]

[0065] The unit direction vectors of the three axes of the body coordinate system in the laser measurement coordinate system are:

[0066]

[0067] Where a, b, and c are the undetermined coefficients of the equation of the plane containing the two-dimensional coplanar target in the laser measurement coordinate system, determined by the least squares method. The results were obtained through fitting.

[0068] Calibrating the relationship between the body coordinate system and the laser measurement coordinate system:

[0069] Let J be the origin of the laser measurement coordinate system. The unit direction vector of the coordinate axes in the laser measurement coordinate system. θ and ψ represent the body coordinate system and the laser measurement coordinate system, respectively. The Euler angles of the axis rotation, taken in a 3-1-2 rotation order, yield the rotation matrix R from the laser measurement coordinate system to the body coordinate system of the reflecting surface:

[0070]

[0071] Therefore, the pose of the body coordinate system relative to the laser measurement coordinate system can be obtained as follows:

[0072]

[0073] Step 2: Install an infrared camera on the microwave payload and calibrate the coordinate transformation link from the infrared camera coordinate system to the star-sensor support coordinate system on the ground.

[0074] The satellite system is defined as follows:

[0075] Origin of coordinate system B For the satellite's center of mass, +x B The axis points in the direction of satellite flight, +z B The axis points to the sub-star point, +y B axis and +x B +z B The axes form a right-handed rectangular coordinate system.

[0076] The coordinate system for the main inversion design is defined as follows:

[0077] Origin of coordinate system M For the geometric vertex of the parabolic surface of revolution, +z M The axis coincides with the focal axis of the paraboloid of revolution and points in the positive normal direction to the geometric vertex of the paraboloid of revolution, +x M Axis perpendicular + z M The axis points to a characteristic direction of the parabolic surface of revolution (with respect to the satellite's body coordinate system +z). B (The axes form an angle of 66.003°), +y M axis and +z M +x M The axes form a right-handed rectangular coordinate system.

[0078] The coordinate system of the infrared camera is defined as follows:

[0079] Origin of coordinate system C Located at the center of the camera's focal plane, +z C The axis represents the direction of the optical axis, +x C axis and +y C The axes are two orthogonal references defined on the camera's focal plane.

[0080] The coordinate system of the star-sensor bracket is defined as follows:

[0081] Origin of coordinate system S Located at the center of the star-sensitive photosensitive surface, +z SThe axis points in the direction of the star sensor's principal optical axis, +x S axis and +y S The axes are two orthogonal references defined on the star-sensitive photosensitive surface.

[0082] On the ground, the three Euler angles (α, β) of the infrared camera coordinate system relative to the master-reverse design coordinate system are first calibrated using a 3-2-3 rotation sequence. C ,β C γ C This yields the rotation matrix from the master inverse design coordinate system to the infrared camera coordinate system. Then, the rotation matrix of the principal anti-design coordinate system relative to the satellite body system is determined using a 3-2-3 rotation sequence. Finally, the rotation matrix of the star-sensor coordinate system relative to the satellite's own system was determined.

[0083] Step 3: Use the common point conversion target to realize the reference transfer of the laser measurement results of the coplanar laser target, calibrate the thermal deformation law of each reflecting surface on the track under the unified reference, and obtain the true pointing relationship of the infrared camera optical axis relative to the reference coordinate system after conversion.

[0084] Specifically, such as Figure 4 As shown, the measurement reference transfer process is as follows:

[0085] Let P = (p1, p2, ..., p n ) and Q = (q1, q2, ..., q n Let be the coordinates of the common measurement target obtained by different measurement methods, then we have:

[0086] q i =Rp i +t(i=1,2,…,n).

[0087] To obtain the best estimates R and t, the singular value decomposition method is used to solve the optimal solution problem:

[0088]

[0089] The process of transferring the measurement reference between laser measurement coordinate system II and laser measurement coordinate system I is as follows:

[0090]

[0091] in, The rotation matrix representing the distance between laser measurement coordinate systems one and two and the common measurement coordinate system is obtained by measuring the common measurement target on a pair of anti-bracing rods using laser scanning measurement devices one and two.

[0092] The process of transferring the measurement datum between the laser measurement coordinate system and the master inverse design coordinate system is as follows:

[0093]

[0094] in, The rotation matrix representing the laser measurement coordinate system to the main inverse design coordinate system is obtained by measuring the common measurement target in the central region of the main inverse surface using a laser scanning measurement device.

[0095] Establish body coordinate systems for each reflecting surface within the master reflector design coordinate system. Calculate the pose relationship between each reflecting surface's body coordinate system and the master reflector design coordinate system using the process described in step 1. Let M be the origin of the master reflector design system. The unit direction vector of the coordinate axis of the master inverse design system. θ and ψ are the body coordinate system around the principal inverse design system, respectively. The Euler angles of the axis rotation, taken in a 3-1-2 rotation order, yield the rotation matrix R from the principal anti-reflection design frame to the body coordinate system of the reflecting surface:

[0096]

[0097] Therefore, the pose of the body coordinate system relative to the principal inverse design system can be obtained as follows:

[0098]

[0099] Includes: a reversed pose: (Q′) F1 , θ′ F1 , ψ′ F1 ), Second reverse pose: (Q′ F2 , θ′ F2 , ψ′ F2 ), Master Reversed Pose: (Q′) F , θ′ F , ψ′ F ).

[0100] Based on the infrared camera optical path model, the actual orientation of the camera's optical axis in the principal reflector design coordinate system under the condition of on-orbit thermal deformation of the reflector antenna is:

[0101]

[0102] Where: l C It is the ideal orientation of the infrared camera's optical axis in the camera coordinate system, which can be calculated from the infrared camera's rear optical path imaging model. f(·) represents the ideal optical path model of the microwave load.

[0103] Step 4: Calculate the three-axis pointing deviation angle of the infrared camera optical axis relative to the star sensor support coordinate system based on the infrared landmark observation results, and then calibrate the overall thermal deformation law of "microwave payload-star sensor" at different landmark observation times in orbit.

[0104] The optical axis pointing deviation angle of the infrared camera is defined as follows:

[0105] The infrared camera optical axis pointing deviation angle ( θ ma , ψ ma This describes the combined effect of thermal deformation of each reflective surface and thermal deformation of the star-sensitive support coordinate system on the optical axis orientation of the infrared camera. Using 3-1-2 as the transformation sequence, the actual orientation of the infrared camera's optical axis in the star-sensitive support coordinate system is:

[0106]

[0107] The infrared landmark images transmitted from satellites are matched with the land-sea boundary landmark dataset in the landmark navigation database to obtain landmark matching results.

[0108] Let P be the j-th image point in the landmark matching result at any landmark observation time ti (1≤i≤N). j (j = 1, 2, ..., n, where n is a positive integer representing the number of landmarks observed in each landmark observation mission) The ideal pointing unit vector corresponding to this in the star-sensor coordinate system is l S，ti，j =[x j y j , z j ] S，ti And will be with image point P j Grid point P′ in the corresponding landmark dataset i The geodetic coordinates are extracted and combined with auxiliary data such as orbit and attitude transmitted from the satellite, then transformed into the star-sensitive support coordinate system at time ti to obtain the image point P. i The corresponding actual pointing unit vector l′ S，ti，j =[x′ i y′ i , z′ i ] S，ti ,l′ S，ti，j and l S，ti，j The difference is the infrared camera optical axis pointing deviation angle at time ti. θ ma , ψ ma ) ti Using a 3-1-2 rotation matrix, it can be written in the following form:

[0109]

[0110] The least squares method was used to calculate the three-axis pointing deviation angles of the infrared camera's optical axis relative to the star-pointing bracket coordinate system at different landmark observation times. θ ma , ψ ma )t i，The optical axis pointing deviation angle of the infrared camera was determined using a multi-order Fourier series model. θ ma , ψ ma The variation pattern over time was fitted as follows:

[0111]

[0112] Where: (A0, C0, E0) are the constant components of the infrared camera's optical axis pointing deviation angle, (A1, ..., B6, C1, ..., D6, E1, ..., F6) are the amplitudes of each periodic component of the infrared camera's optical axis pointing deviation angle, ω is the fundamental frequency of the periodic component (i.e., the Earth's rotational angular velocity), and t is time. Thus, the fitting coefficients are calculated. Then, the true orientation of the infrared camera's optical axis in the star-sensitive support coordinate system at any time t can be obtained.

[0113] Step 5: Calculate the on-orbit thermal deformation law of the star sensor based on the overall thermal deformation law of the "microwave payload-star sensor" and the thermal deformation law of the microwave payload itself, and realize the on-orbit decoupling calibration of the thermal deformation of the "payload-star sensor" of the geostationary microwave detection satellite.

[0114] Infrared landmark observations can identify the time-varying pattern of the three-axis pointing deviation angle of the infrared camera optical axis in the star-sensor support coordinate system caused by the overall thermal deformation of the "microwave payload-star sensor". Therefore, the true time-varying pattern of the infrared camera optical axis pointing in the star-sensor support coordinate system can be obtained as follows:

[0115]

[0116] Laser scanning measurement can identify the thermal deformation patterns of each reflective surface. After conversion, the time-varying pattern of the true pointing of the infrared camera's optical axis in the main reflector design system can be obtained as follows:

[0117]

[0118] remember θ S，t ψ S，t The three-axis attitude deviation of the star-sensor bracket relative to its nominal mounting orientation at time t can be obtained by using a 3-1-2 rotation sequence as follows: pass The actual orientation of the infrared camera's optical axis in the star-sensitive support coordinate system and the main reflector design system can be linked as follows:

[0119]

[0120] The on-orbit thermal deformation pattern of the star sensor itself can be determined from the least squares method. θ S，t , ψS，t During the operation of a geostationary microwave sounding satellite, when special observation requirements result in a lack of infrared landmark observation data, the on-orbit thermal deformation law (Q′) of the calibrated three-reflector solid-surface antenna can be used as a reference. F2 ,…,ψ′ F2 ;Q′ F1 ,…,ψ′ F1 ;Q′ F ,…,ψ′ F ) and the on-orbit thermal deformation law of star sensor ( θ S，t , ψ S，t To ensure the stability and continuity of business operations, real-time image navigation registration is performed in orbit.

[0121] This invention also provides an on-orbit decoupling calibration system for geostationary microwave sounding satellite payload-star-sensor thermal deformation. This system can be implemented by executing the steps of the on-orbit decoupling calibration method for geostationary microwave sounding satellite payload-star-sensor thermal deformation. In other words, those skilled in the art can understand the on-orbit decoupling calibration method for geostationary microwave sounding satellite payload-star-sensor thermal deformation as a preferred embodiment of the on-orbit decoupling calibration system for geostationary microwave sounding satellite payload-star-sensor thermal deformation.

[0122] Those skilled in the art will understand that, besides implementing the system and its various devices, modules, and units provided by this invention in the form of purely computer-readable program code, the same functions can be achieved entirely through logical programming of the method steps, enabling the system and its various devices, modules, and units to function in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers. Therefore, the system and its various devices, modules, and units provided by this invention can be considered a hardware component, and the devices, modules, and units included therein for implementing various functions can also be considered structures within the hardware component; alternatively, the devices, modules, and units for implementing various functions can be considered both software modules implementing the method and structures within the hardware component.

[0123] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0124] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A geostationary orbit microwave detection satellite payload-star-sensitive thermal deformation on-orbit decoupling calibration device, characterized in that, include: Microwave payload, star sensor, laser scanning measurement camera, common point conversion target, coplanar laser target and infrared camera; The microwave payload includes: an antenna main reflector, a first secondary reflector, a second secondary reflector, an antenna strut, and a quasi-optical system; The antenna's main reflector, first secondary reflector, and second secondary reflector are mounted on the top plate of the satellite platform via antenna struts; the quasi-optical system is mounted and embedded in the satellite platform via a quasi-optical frame. The microwave payload and the satellite platform are installed at an angle according to preset requirements; The star sensor is installed on the side plate of the geostationary microwave detection satellite platform to ensure that the satellite pointing reference is consistent with the attitude control reference. The laser scanning measurement cameras are respectively mounted on a pair of anti-bracing rods and below the collimation frame; The infrared camera and the quasi-optical frame are mounted together with a common reference. The common point conversion target is affixed to the central area of ​​a pair of anti-bracing rods and the main anti-shape surface; The coplanar laser targets are respectively arranged on three reflective surfaces; The laser scanning measurement camera includes two sets. The laser scanning measurement camera installed on the first set of anti-support rods monitors the on-orbit thermal deformation of the main anti-reflector by measuring the spatial position of the target points on the main anti-reflector. The laser scanning measurement camera installed below the quasi-optical frame monitors the on-orbit thermal deformation of the first and second anti-reflectors by measuring the spatial position of the target points on the first and second anti-reflectors. The common point conversion target is respectively arranged in the central area of ​​a pair of anti-bracing rods and the main anti-plane. The common point conversion target is used to realize the reference transfer of the laser measurement results of the coplanar laser target of each reflecting surface, and to unify the measurement data of the two sets of laser scanning measurement cameras under the main anti-plane design coordinate system. The distance-angle spherical coordinates measured by the laser scanning measurement camera in the laser measurement coordinate system are converted into Cartesian three-dimensional space coordinates, and the on-orbit thermal deformation is monitored based on the Cartesian three-dimensional space coordinates. The infrared camera selects a working band that meets preset requirements to acquire infrared landmark images of cloudless or less cloudy areas. Based on the infrared landmark observation results, the three-axis pointing deviation angle of the infrared camera's optical axis relative to the star sensor support coordinate system is calculated, and then the overall thermal deformation law of the microwave payload-star sensor at different landmark observation times is calibrated on orbit.

2. The on-orbit decoupling calibration device for geostationary microwave detection satellite payload-star-sensitive thermal deformation according to claim 1, characterized in that, The coplanar laser targets are arranged on three reflective surfaces to establish a body coordinate system for each reflective surface. By directly comparing the on-orbit calibration results of the body coordinate system with the ground calibration results, the on-orbit thermal deformation of each reflective surface can be obtained.

3. The on-orbit decoupling calibration device for geostationary microwave detection satellite payload-star-sensitive thermal deformation according to claim 1, characterized in that, The infrared camera uses an antenna-converging lens-rear optical path assembly. The converging lens is mounted on a support rod near the second secondary mirror, and the rear optical path is mounted on the collimating frame and the connecting rod of the second secondary mirror, sharing a common reference with the collimating frame.

4. The on-orbit decoupling calibration device for geostationary microwave detection satellite payload-star-sensitive thermal deformation according to claim 1, characterized in that, The microwave payload's primary reflector is a 5m diameter parabolic rotating surface, the secondary reflector is a 1m diameter hyperboloid rotating surface, and the third secondary reflector is a 1.4m diameter planar surface.

5. The on-orbit decoupling calibration device for geostationary microwave detection satellite payload-star-sensitive thermal deformation according to claim 1, characterized in that, After the microwave load is installed at an angle, the main inverse design coordinate system's line of sight to the ground... With satellite body coordinate system Shaft Angle.

6. The on-orbit decoupling calibration device for geostationary microwave detection satellite payload-star-sensitive thermal deformation according to claim 1, characterized in that, The star sensor is mounted on a star sensor bracket, and the star sensor bracket and the satellite platform are integrated into a mechanical and thermal structure.

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