Multi-pose measuring device relative installation error calibration and compensation method and system

Abstract

The application provides a multi-pose measuring device relative installation error calibration and compensation method and system, and relates to the technical field of spaceflight. The application comprises the following steps: S1, measuring a star sensor pose and converting coordinates; S2, calibrating high-precision inertial navigation and star sensor relative installation errors; S3, calibrating high-precision inertial navigation and low-precision inertial navigation self-alignment relative installation errors; and S4, compensating for the relative installation errors. The application calibrates and compensates for the relative installation errors among the high-precision inertial measurement combination, the low-precision inertial measurement combination and the star sensor by using the inertial navigation static self-alignment and the star sensor static pose measurement in the static state, so as to realize the consistency of the space pose reference in the dynamic multi-source combined navigation information fusion.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and to a method for rapid calibration and compensation of relative installation errors of inertial navigation and star-aware multi-attitude measurement equipment. Specifically, it relates to a method for rapid calibration and compensation of relative installation errors of multiple different types of attitude measurement equipment used in multi-source integrated navigation through attitude measurement algorithms. In particular, it relates to a method and system for calibration and compensation of relative installation errors of multi-attitude measurement equipment. Background Technology

[0002] Currently, multi-source integrated navigation has been widely used in various life scenarios such as aerospace, aviation, autonomous driving, underwater navigation, and indoor positioning. When integrating multiple sources, it is often necessary to fuse the attitude measurements of various sensors (high-precision and low-precision inertial navigation, star sensors, and magnetic compasses). However, the installation positions and installation references of each sensor are usually inconsistent, and the measured attitudes cannot be directly fused. Therefore, before multi-source integrated navigation, it is necessary to calibrate and compensate for the relative installation errors of different attitude measurement sensors.

[0003] Traditional calibration methods often involve installing prisms or cubic mirrors on various sensors and measuring the relative installation angle error using a theodolite or photoelectric autocollimator. This method requires pre-calibrating the relative error between the prisms or cubic mirrors and the measurement coordinate system. Furthermore, the operation is cumbersome and cannot be quickly calibrated and compensated. Summary of the Invention

[0004] To address the shortcomings of existing technologies, this invention provides a method and system for calibrating and compensating for relative installation errors of multi-pose measurement equipment.

[0005] According to the present invention, a method and system for calibrating and compensating relative installation errors of a multi-pose measuring device are provided, the solution of which is as follows:

[0006] Firstly, a method for calibrating and compensating relative installation errors of a multi-pose measurement device is provided, the method comprising:

[0007] Step S1: Measure the star sensor attitude and transform the coordinates;

[0008] Step S2: Calibrate the relative installation error of the high-precision inertial navigation system and star sensor;

[0009] Step S3: Calibrate the relative installation error of the self-alignment of the high-precision inertial navigation system and the low-precision inertial navigation system;

[0010] Step S4: Compensate for relative installation errors.

[0011] Preferably, step S1 includes:

[0012] With the carrier platform's X-axis pointing vertically upwards, the star sensor performs attitude measurements at night with its orientation also pointing vertically upwards, obtaining the attitude quaternions from the star sensor's intrinsic frame to the J2000 inertial frame.

[0013] Based on the obtained attitude transformation from the star-sensitive datum to the J2000 inertial frame, the attitude array of the star-sensitive datum to the North-East coordinate system is converted. The specific conversion process is as follows:

[0014] First, from the star-aware system to the attitude quaternion of the J2000 inertial frame. The attitude array is constructed as follows:

[0015]

[0016] In the formula, q0, q1, q2, and q3 are the four elements of the attitude quaternion from the star-sensor body system to the J2000 inertial frame;

[0017] Secondly, the transformation matrix from the J2000 inertial frame to the Earth coordinate system is:

[0018]

[0019] In the formula, A is the transformation matrix between the quasi-Earth-fixed coordinate system and the Earth coordinate system, B is the Earth rotation transformation matrix, C is the nutation transformation matrix, and D is the precession transformation matrix;

[0020] Finally, the process of converting the attitude array from the star-sensitive intrinsic system to the J2000 inertial frame to the attitude array from the star-sensitive intrinsic system to the North-East coordinate system is as follows:

[0021]

[0022] in,

[0023] In the formula, The Earth-to-navigation system conversion matrix, where λ is the carrier's longitude and L is the carrier's latitude.

[0024] Preferably, step S2 includes:

[0025] The X-axis of the carrier platform is vertically upward, as are the X3 axis of the star sensor and the X1 axis of the high-precision inertial navigation system.

[0026] The star sensor completes attitude measurement and acquires the star sensor attitude array from the local system to the north-sky-east coordinate system according to step S1.

[0027] The high-precision inertial navigation system (INS) performs self-alignment based on the bound geographic parameters. After alignment, the attitude quaternions of the INS relative to the North-Sky-East coordinate system are obtained, and the attitude matrix is ​​obtained accordingly.

[0028]

[0029] In the formula, q 10 q 11 q 12 q 13 These are the four elements of the attitude quaternion of the high-precision inertial navigation system relative to the North-East coordinate system;

[0030] Obtain the relative installation error matrix between the high-precision inertial navigation system X1Y1Z1 and the star sensor system X3Y3Z3:

[0031]

[0032] Preferably, step S3 includes: placing the carrier platform horizontally with the high-precision inertial navigation Y1 axis and the low-precision inertial navigation Y2 axis pointing vertically upwards, and self-aligning to obtain the high-precision inertial navigation pitch angle θ. h Roll angle γ h and low-precision inertial navigation pitch angle θ l Roll angle γ l ;

[0033] The carrier platform is rotated 90° and placed horizontally, with the high-precision inertial navigation Z1 axis and the low-precision inertial navigation Z2 axis both pointing vertically upwards. The high-precision inertial navigation pitch angle θ is obtained through self-alignment. h 'and low-precision inertial navigation pitch angle θ l ';

[0034] When the carrier platform is rotated 90° and placed horizontally, the high-precision and low-precision inertial navigation pitch angles are the same as the high-precision and low-precision inertial navigation yaw angles when the carrier platform is normally placed horizontally, that is:

[0035] ψ h =θ h '

[0036] ψ l =θ l '

[0037] Calculate the attitude matrix of the high- and low-precision inertial navigation system relative to the North-East coordinate system based on the pitch angle, roll angle, and yaw angle of the high- and low-precision inertial navigation systems:

[0038]

[0039]

[0040] Based on the attitude matrices of the high-precision and low-precision inertial navigation systems relative to the North-East coordinate system, the relative installation error matrices of the high-precision inertial navigation system X1Y1Z1 and the low-precision inertial navigation system X2Y2Z2 are calculated:

[0041]

[0042] Preferably, step S4 includes:

[0043] Multi-source integrated navigation systems use a high-precision inertial navigation body coordinate system as the coordinate reference;

[0044] Through the relative installation error matrix Compensation is used to adjust the attitude matrix measured dynamically in real time by the low-precision inertial navigation system. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0045]

[0046] In the formula: To compensate for low-precision inertial navigation real-time dynamic measurement attitude matrix after installation error;

[0047] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0048] Through the relative installation error matrix Compensation is applied to the attitude matrix measured dynamically in real time by the star sensor. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0049]

[0050] In the formula: The attitude matrix of the star sensor is dynamically measured in real time to compensate for installation errors;

[0051] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0052] Secondly, a system for calibrating and compensating relative installation errors of a multi-pose measurement device is provided, the system comprising:

[0053] Module M1: Measures the star sensor attitude and transforms coordinates;

[0054] Module M2: Calibrates the relative installation error of high-precision inertial navigation and star sensors;

[0055] Module M3: Calibrates the relative installation error of high-precision and low-precision inertial navigation systems for self-alignment;

[0056] Module M4: Compensates for relative installation errors.

[0057] Preferably, the module M1 includes:

[0058] With the carrier platform's X-axis pointing vertically upwards, the star sensor performs attitude measurements at night with its orientation also pointing vertically upwards, obtaining the attitude quaternions from the star sensor's intrinsic frame to the J2000 inertial frame.

[0059] Based on the obtained attitude transformation from the star-sensitive datum to the J2000 inertial frame, the attitude array of the star-sensitive datum to the North-East coordinate system is converted. The specific conversion process is as follows:

[0060] First, from the star-aware system to the attitude quaternion of the J2000 inertial frame. The attitude array is constructed as follows:

[0061]

[0062] In the formula, q0, q1, q2, and q3 are the four elements of the attitude quaternion from the star-sensor body system to the J2000 inertial frame;

[0063] Secondly, the transformation matrix from the J2000 inertial frame to the Earth coordinate system is:

[0064]

[0065] In the formula, A is the transformation matrix between the quasi-Earth-fixed coordinate system and the Earth coordinate system, B is the Earth rotation transformation matrix, C is the nutation transformation matrix, and D is the precession transformation matrix;

[0066] Finally, the process of converting the attitude array from the star-sensitive intrinsic system to the J2000 inertial frame to the attitude array from the star-sensitive intrinsic system to the North-East coordinate system is as follows:

[0067]

[0068] in,

[0069] In the formula, The Earth-to-navigation system conversion matrix, where λ is the carrier's longitude and L is the carrier's latitude.

[0070] Preferably, the module M2 includes:

[0071] The X-axis of the carrier platform is vertically upward, as are the X3 axis of the star sensor and the X1 axis of the high-precision inertial navigation system.

[0072] The star sensor completes attitude measurement and acquires the star sensor attitude array from the local system to the north-sky-east coordinate system according to step S1.

[0073] The high-precision inertial navigation system (INS) performs self-alignment based on the bound geographic parameters. After alignment, the attitude quaternions of the INS relative to the North-Sky-East coordinate system are obtained, and the attitude matrix is ​​obtained accordingly.

[0074]

[0075] In the formula, q 10 q 11 q 12 q 13 These are the four elements of the attitude quaternion of the high-precision inertial navigation system relative to the North-East coordinate system;

[0076] Obtain the relative installation error matrix between the high-precision inertial navigation system X1Y1Z1 and the star sensor system X3Y3Z3:

[0077]

[0078] Preferably, the module M3 includes: a carrier platform placed horizontally, with the high-precision inertial navigation Y1 axis and the low-precision inertial navigation Y2 axis both pointing vertically upwards, and self-aligned to obtain the high-precision inertial navigation pitch angle θ. h Roll angle γ h and low-precision inertial navigation pitch angle θ l Roll angle γ l ;

[0079] The carrier platform is rotated 90° and placed horizontally, with the high-precision inertial navigation Z1 axis and the low-precision inertial navigation Z2 axis both pointing vertically upwards. The high-precision inertial navigation pitch angle θ is obtained through self-alignment. h 'and low-precision inertial navigation pitch angle θ l ';

[0080] When the carrier platform is rotated 90° and placed horizontally, the high-precision and low-precision inertial navigation pitch angles are the same as the high-precision and low-precision inertial navigation yaw angles when the carrier platform is normally placed horizontally, that is:

[0081] ψ h =θ h '

[0082] ψ l =θ l '

[0083] Calculate the attitude matrix of the high- and low-precision inertial navigation system relative to the North-East coordinate system based on the pitch angle, roll angle, and yaw angle of the high- and low-precision inertial navigation systems:

[0084]

[0085]

[0086] Based on the attitude matrices of the high-precision and low-precision inertial navigation systems relative to the North-East coordinate system, the relative installation error matrices of the high-precision inertial navigation system X1Y1Z1 and the low-precision inertial navigation system X2Y2Z2 are calculated:

[0087]

[0088] Preferably, the module M4 includes:

[0089] Multi-source integrated navigation systems use a high-precision inertial navigation body coordinate system as the coordinate reference;

[0090] Through the relative installation error matrix Compensation is used to adjust the attitude matrix measured dynamically in real time by the low-precision inertial navigation system. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0091]

[0092] In the formula: To compensate for low-precision inertial navigation real-time dynamic measurement attitude matrix after installation error;

[0093] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0094] Through the relative installation error matrix Compensation is applied to the attitude matrix measured dynamically in real time by the star sensor. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0095]

[0096] In the formula: The attitude matrix of the star sensor is dynamically measured in real time to compensate for installation errors;

[0097] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0098] Compared with the prior art, the present invention has the following beneficial effects:

[0099] 1. Utilizing static self-alignment of inertial navigation, the static attitude measurement of star sensors calibrates and compensates for the relative installation errors between high-precision inertial measurement systems, low-precision inertial measurement systems, and star sensors in a static state, so as to achieve consistency of spatial attitude reference when multi-source integrated navigation information is fused in a dynamic state.

[0100] 2. This method avoids the cumbersome operation of traditional prism or cubic mirror calibration methods and can achieve rapid calibration and compensation.

[0101] Other beneficial effects of the present invention will be explained in detail through the introduction of specific technical features and technical solutions in specific embodiments. Those skilled in the art should be able to understand the beneficial technical effects brought about by these technical features and technical solutions through the introduction of these technical features and technical solutions. Attached Figure Description

[0102] 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:

[0103] Figure 1 This is a schematic diagram of the installation relationship and coordinate system of the inertial navigation system and star sensor on the carrier platform according to an embodiment of the present invention;

[0104] Figure 2 These are the implementation steps of an embodiment of the present invention. Detailed Implementation

[0105] 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.

[0106] This invention provides a method for calibrating and compensating for relative installation errors of multi-attitude measurement devices. It rapidly calibrates and compensates for the relative installation errors of three attitude measurement devices: a high-precision inertial navigation system (INS), a low-precision INS, and a star sensor. The installation relationship and coordinate system of the three devices on the carrier platform are shown below. Figure 1 As shown, the method specifically includes the following, see reference. Figure 2 As shown:

[0107] Step S1: Measure the star sensor attitude and transform the coordinates.

[0108] With the carrier platform's X-axis pointing vertically upwards, the star sensor performs attitude measurements at night with its orientation also pointing vertically upwards, obtaining the attitude quaternions from the star sensor's intrinsic frame to the J2000 inertial frame.

[0109] Based on the obtained attitude transformation from the star-sensitive datum to the J2000 inertial frame, the attitude array of the star-sensitive datum to the North-East coordinate system is converted. The specific conversion process is as follows:

[0110] First, from the star-aware system to the attitude quaternion of the J2000 inertial frame. The attitude array is constructed as follows:

[0111]

[0112] In the formula, q0, q1, q2, and q3 are the four elements of the attitude quaternion from the star-sensor body system to the J2000 inertial frame;

[0113] Secondly, the transformation matrix from the J2000 inertial frame to the Earth coordinate system is:

[0114]

[0115] In the formula, A is the transformation matrix between the quasi-Earth-fixed coordinate system and the Earth coordinate system, B is the Earth rotation transformation matrix, C is the nutation transformation matrix, and D is the precession transformation matrix;

[0116] Finally, the process of converting the attitude array from the star-sensitive intrinsic system to the J2000 inertial frame to the attitude array from the star-sensitive intrinsic system to the North-East coordinate system is as follows:

[0117]

[0118] in,

[0119] In the formula, The Earth-to-navigation system conversion matrix, where λ is the carrier's longitude and L is the carrier's latitude.

[0120] Step S2: Calibrate the relative installation error of the high-precision inertial navigation system and star sensor.

[0121] This step specifically includes: the X-axis of the carrier platform is vertically oriented towards the sky, and the X3 axis of the star sensor and the X1 axis of the high-precision inertial navigation are both vertically oriented towards the sky;

[0122] The star sensor completes attitude measurement and acquires the star sensor attitude array from the local system to the north-sky-east coordinate system according to step S1.

[0123] The high-precision inertial navigation system (INS) performs self-alignment based on the bound geographical parameters (latitude, longitude, altitude, gravity). After alignment, the attitude quaternions of the INS relative to the North-East coordinate system are obtained, and the attitude matrix is ​​obtained accordingly.

[0124]

[0125] In the formula, q 10 q 11 q 12 q 13 These are the four elements of the attitude quaternion of the high-precision inertial navigation system relative to the North-East coordinate system;

[0126] Obtain the relative installation error matrix between the high-precision inertial navigation system X1Y1Z1 and the star sensor system X3Y3Z3:

[0127]

[0128] Step S3: Calibrate the relative installation error of the self-alignment of the high-precision inertial navigation system and the low-precision inertial navigation system;

[0129] Step S3 specifically includes: placing the carrier platform horizontally with the high-precision inertial navigation Y1 axis and the low-precision inertial navigation Y2 axis both pointing vertically upwards, and self-aligning to obtain the high-precision inertial navigation pitch angle θ. hRoll angle γ h and low-precision inertial navigation pitch angle θ l Roll angle γ l ;

[0130] The carrier platform is rotated 90° and placed horizontally, with the high-precision inertial navigation Z1 axis and the low-precision inertial navigation Z2 axis both pointing vertically upwards. The high-precision inertial navigation pitch angle θ' is obtained through self-alignment. h and low-precision inertial navigation pitch angle θ' l ;

[0131] When the carrier platform is rotated 90° and placed horizontally, the high-precision and low-precision inertial navigation pitch angles are the same as the high-precision and low-precision inertial navigation yaw angles when the carrier platform is normally placed horizontally, that is:

[0132] ψ h =θ' h

[0133] ψ l =θ' l

[0134] Calculate the attitude matrix of the high- and low-precision inertial navigation system relative to the North-East coordinate system based on the pitch angle, roll angle, and yaw angle of the high- and low-precision inertial navigation systems:

[0135]

[0136]

[0137] Based on the attitude matrices of the high-precision and low-precision inertial navigation systems relative to the North-East coordinate system, the relative installation error matrices of the high-precision inertial navigation system X1Y1Z1 and the low-precision inertial navigation system X2Y2Z2 are calculated:

[0138]

[0139] Step S4: Compensate for relative installation errors.

[0140] Step S4 specifically includes:

[0141] Multi-source integrated navigation systems use a high-precision inertial navigation body coordinate system as the coordinate reference;

[0142] Through the relative installation error matrix Compensation is used to adjust the attitude matrix measured dynamically in real time by the low-precision inertial navigation system. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0143]

[0144] In the formula: To compensate for low-precision inertial navigation real-time dynamic measurement attitude matrix after installation error;

[0145] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0146] Through the relative installation error matrix Compensation is applied to the attitude matrix measured dynamically in real time by the star sensor. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0147]

[0148] In the formula: The attitude matrix of the star sensor is dynamically measured in real time to compensate for installation errors;

[0149] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0150] The present invention also provides a system for calibrating and compensating the relative installation error of a multi-position measuring device. The system can be implemented by executing the process steps of the method for calibrating and compensating the relative installation error of a multi-position measuring device. That is, those skilled in the art can understand the method for calibrating and compensating the relative installation error of a multi-position measuring device as a preferred embodiment of the system for calibrating and compensating the relative installation error of a multi-position measuring device.

[0151] Module M1: Measures the attitude of the star sensor and transforms its coordinates.

[0152] With the carrier platform's X-axis pointing vertically upwards, the star sensor performs attitude measurements at night with its orientation also pointing vertically upwards, obtaining the attitude quaternions from the star sensor's intrinsic frame to the J2000 inertial frame.

[0153] Based on the obtained attitude transformation from the star-sensitive datum to the J2000 inertial frame, the attitude array of the star-sensitive datum to the North-East coordinate system is converted. The specific conversion process is as follows:

[0154] First, from the star-aware system to the attitude quaternion of the J2000 inertial frame. The attitude array is constructed as follows:

[0155]

[0156] In the formula, q0, q1, q2, and q3 are the four elements of the attitude quaternion from the star-sensor body system to the J2000 inertial frame;

[0157] Secondly, the transformation matrix from the J2000 inertial frame to the Earth coordinate system is:

[0158]

[0159] In the formula, A is the transformation matrix between the quasi-Earth-fixed coordinate system and the Earth coordinate system, B is the Earth rotation transformation matrix, C is the nutation transformation matrix, and D is the precession transformation matrix;

[0160] Finally, the process of converting the attitude array from the star-sensitive intrinsic system to the J2000 inertial frame to the attitude array from the star-sensitive intrinsic system to the North-East coordinate system is as follows:

[0161]

[0162] in,

[0163] In the formula, The Earth-to-navigation system conversion matrix, where λ is the carrier's longitude and L is the carrier's latitude.

[0164] Module M2: Calibrates the relative installation error of high-precision inertial navigation and star sensors.

[0165] This module specifically includes: the carrier platform with its X-axis pointing vertically upwards, and the star sensor's X3 axis and the high-precision inertial navigation X1 axis both pointing vertically upwards;

[0166] The star sensor completes attitude measurement according to module M1 and acquires the star sensor attitude array from the local system to the north-sky-east coordinate system.

[0167] The high-precision inertial navigation system (INS) performs self-alignment based on the bound geographical parameters (latitude, longitude, altitude, gravity). After alignment, the attitude quaternions of the INS relative to the North-East coordinate system are obtained, and the attitude matrix is ​​obtained accordingly.

[0168]

[0169] In the formula, q 10 q 11 q 12 q 13 These are the four elements of the attitude quaternion of the high-precision inertial navigation system relative to the North-East coordinate system;

[0170] Obtain the relative installation error matrix between the high-precision inertial navigation system X1Y1Z1 and the star sensor system X3Y3Z3:

[0171]

[0172] Module M3: Calibrates the relative installation error of high-precision and low-precision inertial navigation systems for self-alignment;

[0173] Module M3 specifically includes: a carrier platform placed horizontally, with both the high-precision inertial navigation Y1 axis and the low-precision inertial navigation Y2 axis pointing vertically upwards, and self-aligning to obtain the high-precision inertial navigation pitch angle θ. h Roll angle γ h and low-precision inertial navigation pitch angle θ l Roll angle γ l ;

[0174] The carrier platform is rotated 90° and placed horizontally, with the high-precision inertial navigation Z1 axis and the low-precision inertial navigation Z2 axis both pointing vertically upwards. The high-precision inertial navigation pitch angle θ' is obtained through self-alignment. h and low-precision inertial navigation pitch angle θ' l ;

[0175] When the carrier platform is rotated 90° and placed horizontally, the high-precision and low-precision inertial navigation pitch angles are the same as the high-precision and low-precision inertial navigation yaw angles when the carrier platform is normally placed horizontally, that is:

[0176] ψ h =θ' h

[0177] ψ l =θ' l

[0178] Calculate the attitude matrix of the high- and low-precision inertial navigation system relative to the North-East coordinate system based on the pitch angle, roll angle, and yaw angle of the high- and low-precision inertial navigation systems:

[0179]

[0180]

[0181] Based on the attitude matrices of the high-precision and low-precision inertial navigation systems relative to the North-East coordinate system, the relative installation error matrices of the high-precision inertial navigation system X1Y1Z1 and the low-precision inertial navigation system X2Y2Z2 are calculated:

[0182]

[0183] Module M4: Compensates for relative installation errors.

[0184] The M4 module specifically includes:

[0185] Multi-source integrated navigation systems use a high-precision inertial navigation body coordinate system as the coordinate reference;

[0186] Through the relative installation error matrix Compensation is used to adjust the attitude matrix measured dynamically in real time by the low-precision inertial navigation system. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0187]

[0188] In the formula: To compensate for low-precision inertial navigation real-time dynamic measurement attitude matrix after installation error;

[0189] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0190] Through the relative installation error matrix Compensation is applied to the attitude matrix measured dynamically in real time by the star sensor. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0191]

[0192] In the formula: The attitude matrix of the star sensor is dynamically measured in real time to compensate for installation errors;

[0193] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0194] The present invention will now be described in more detail.

[0195] This invention provides a method for calibrating and compensating relative installation errors of a multi-attitude measurement device. Utilizing static self-alignment of inertial navigation and static attitude measurement of star sensors, the method calibrates and compensates for relative installation errors between high-precision and low-precision inertial measurement systems (INS) and star sensors in a static state, thereby achieving consistency of spatial attitude reference during dynamic multi-source integrated navigation information fusion. This method avoids the cumbersome operation of traditional prism or cubic mirror calibration methods and enables rapid calibration and compensation.

[0196] Figure 1 This is a schematic diagram of the installation relationship and coordinate system of the inertial navigation system and star sensor on the carrier platform according to an embodiment of the present invention, including:

[0197] Coordinate system X1Y1Z1: High-precision inertial navigation body coordinate system;

[0198] Coordinate system X2Y2Z2: Low-precision inertial navigation body coordinate system;

[0199] Coordinate system X3Y3Z3: Star sensor body coordinate system.

[0200] Figure 2 The implementation steps of a method for rapid calibration and compensation of relative installation errors of inertial navigation and star-sensor multi-attitude measurement equipment according to an embodiment of the present invention include:

[0201] Step S1: Star sensor attitude measurement and coordinate transformation.

[0202] With the carrier platform's X-axis pointing vertically upwards, the star sensor performs attitude measurements at night with its orientation also pointing vertically upwards, obtaining the attitude quaternions from the star sensor's intrinsic frame to the J2000 inertial frame.

[0203] Based on the obtained attitude transformation from the star-sensitive datum to the J2000 inertial frame, the attitude array of the star-sensitive datum to the North-East coordinate system is converted. The specific conversion process is as follows.

[0204] First, from the star-sensor system to the attitude quaternions of the J2000 inertial frame. The attitude array is constructed as follows:

[0205]

[0206] In the formula, q0, q1, q2, and q3 are the four elements of the attitude quaternion from the star-sensor body system to the J2000 inertial frame.

[0207] The transformation matrix from the J2000 inertial frame to the Earth coordinate system is:

[0208]

[0209] In the formula, A is the transformation matrix between the quasi-Earth-fixed coordinate system and the Earth coordinate system, B is the Earth rotation transformation matrix, C is the nutation transformation matrix, and D is the precession transformation matrix.

[0210] The conversion process from the star-sensitive intrinsic system to the J2000 inertial frame attitude array, and then to the star-sensitive intrinsic system to the North-Eastern coordinate system attitude array, is as follows:

[0211]

[0212] in,

[0213] In the formula, The Earth-to-navigation system conversion matrix, where λ is the carrier's longitude and L is the carrier's latitude.

[0214] Step S2: Calibration of relative installation error between high-precision inertial navigation system and star sensor.

[0215] The carrier platform's X-axis faces vertically upwards, as do the star sensor's X3 axis and the high-precision inertial navigation's X1 axis.

[0216] The star sensor completes attitude measurement and acquires the star sensor attitude array from the local system to the north-sky-east coordinate system according to step S1.

[0217] The high-precision inertial navigation system (INS) performs self-alignment based on the bound geographical parameters (latitude, longitude, altitude, gravity). After alignment, the attitude quaternions of the INS relative to the North-East coordinate system are obtained, and the attitude matrix is ​​obtained accordingly.

[0218]

[0219] In the formula, q 10 q 11 q 12 q 13 These are the four elements of the attitude quaternion of the high-precision inertial navigation system relative to the North-East coordinate system.

[0220] Obtain the relative installation error matrix between the high-precision inertial navigation system X1Y1Z1 and the star sensor system X3Y3Z3:

[0221]

[0222] Step S3: Calibration of relative installation error for high- and low-precision inertial navigation self-alignment

[0223] The carrier platform is placed horizontally with both the high-precision inertial navigation Y1 axis and the low-precision inertial navigation Y2 axis pointing vertically upwards. The high-precision inertial navigation pitch angle θ is obtained through self-alignment. h Roll angle γ h and low-precision inertial navigation pitch angle θ l Roll angle γ l .

[0224] The carrier platform is rotated 90° and placed horizontally, with the high-precision inertial navigation Z1 axis and the low-precision inertial navigation Z2 axis both pointing vertically upwards. The high-precision inertial navigation pitch angle θ' is obtained through self-alignment. h and low-precision inertial navigation pitch angle θ' l .

[0225] When the carrier platform is rotated 90° and placed horizontally, the high-precision and low-precision inertial navigation pitch angles are the same as the high-precision and low-precision inertial navigation yaw angles when the carrier platform is normally placed horizontally, that is:

[0226] ψ h =θ' h

[0227] ψ l =θ' l

[0228] Calculate the attitude matrix of the high- and low-precision inertial navigation system relative to the North-East coordinate system based on the pitch angle, roll angle, and yaw angle of the high- and low-precision inertial navigation systems:

[0229]

[0230]

[0231] Based on the attitude matrices of the high-precision and low-precision inertial navigation systems relative to the North-East coordinate system, the relative installation error matrices of the high-precision inertial navigation system X1Y1Z1 and the low-precision inertial navigation system X2Y2Z2 are calculated:

[0232]

[0233] Step S4: Compensation for relative installation errors.

[0234] Multi-source integrated navigation systems use a high-precision inertial navigation body coordinate system as the coordinate reference.

[0235] Through the relative installation error matrix Compensation is used to adjust the attitude matrix measured dynamically in real time by the low-precision inertial navigation system. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0236]

[0237] In the formula: To compensate for low-precision inertial navigation real-time dynamic measurement attitude matrix after installation error.

[0238] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0239] Through the relative installation error matrix Compensation is applied to the attitude matrix measured dynamically in real time by the star sensor. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system:

[0240]

[0241] In the formula: The attitude matrix of the star sensor is dynamically measured in real time to compensate for installation errors.

[0242] From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

[0243] This invention provides a method and system for calibrating and compensating relative installation errors of a multi-attitude measurement device. By utilizing static self-alignment of inertial navigation and static attitude measurement of star sensors, the relative installation errors between high-precision inertial measurement combination, low-precision inertial measurement combination and star sensors are calibrated and compensated in a static state, so as to achieve consistency of spatial attitude reference when multi-source integrated navigation information is fused in a dynamic state.

[0244] 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, making the system and its various devices, modules, and units of this invention 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 as a hardware component, and the devices, modules, and units included therein for implementing various functions can also be considered as structures within the hardware component; alternatively, the devices, modules, and units for implementing various functions can be considered as both software modules implementing the method and structures within the hardware component.

[0245] 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 method for calibrating and compensating for relative installation errors of a multi-pose measuring device, characterized in that, include: Step S1: Measure the star sensor attitude and transform the coordinates; Step S2: Calibrate the relative installation error of the high-precision inertial navigation system and star sensor; Step S3: Calibrate the relative installation error of the self-alignment of the high-precision inertial navigation system and the low-precision inertial navigation system; Step S4: Compensate for relative installation errors; Step S1 includes: With the carrier platform's X-axis pointing vertically upwards, the star sensor performs attitude measurements at night with its orientation also pointing vertically upwards, obtaining the attitude quaternions from the star sensor's intrinsic frame to the J2000 inertial frame. ; Based on the obtained attitude transformation from the star-sensitive datum to the J2000 inertial frame, the attitude array of the star-sensitive datum to the North-East coordinate system is converted. The specific conversion process is as follows: First, from the star-aware system to the attitude quaternion of the J2000 inertial frame. The attitude array is constructed as follows: In the formula, , , , These are the four elements of the attitude quaternion from the star-aware system to the J2000 inertial frame; Secondly, the transformation matrix from the J2000 inertial frame to the Earth coordinate system is: In the formula, A The transformation matrix between the quasi-Earth-fixed coordinate system and the Earth coordinate system. B This is the Earth's rotation transformation matrix. C Nucleation transformation matrix, D This is the precession transformation matrix; Finally, the process of converting the attitude array from the star-sensitive intrinsic system to the J2000 inertial frame to the attitude array from the star-sensitive intrinsic system to the North-East coordinate system is as follows: in, In the formula, Earth-to-navigation-system conversion matrix, For the longitude of the carrier, L The latitude of the carrier; Step S2 includes: The carrier platform's X-axis is vertically upward, as are the star sensor's X3 axis and the high-precision inertial navigation's X1 axis. The star sensor completes attitude measurement and acquires the star sensor attitude array from the local system to the north-sky-east coordinate system according to step S1. ; The high-precision inertial navigation system (INS) performs self-alignment based on the bound geographic parameters. After alignment, the attitude quaternions of the INS relative to the North-Sky-East coordinate system are obtained, and the attitude matrix is ​​obtained accordingly. In the formula, , , , These are the four elements of the attitude quaternion of the high-precision inertial navigation system relative to the North-East coordinate system; Obtain the relative installation error matrix between the high-precision inertial navigation system X1Y1Z1 and the star sensor system X3Y3Z3: 。 2. The method for calibrating and compensating relative installation errors of a multi-pose measuring device according to claim 1, characterized in that, Step S3 includes: placing the carrier platform horizontally with the high-precision inertial navigation Y1 axis and the low-precision inertial navigation Y2 axis both pointing vertically upwards, and self-aligning to obtain the high-precision inertial navigation pitch angle. Roll angle and low-precision inertial navigation pitch angle Roll angle ; The carrier platform is rotated 90° and placed horizontally, with the high-precision inertial navigation Z1 axis and the low-precision inertial navigation Z2 axis both pointing vertically upwards, and self-aligned to obtain the high-precision inertial navigation pitch angle. and low-precision inertial navigation pitch angle ; When the carrier platform is rotated 90° and placed horizontally, the high-precision and low-precision inertial navigation pitch angles are the same as the high-precision and low-precision inertial navigation yaw angles when the carrier platform is normally placed horizontally, that is: Calculate the attitude matrix of the high- and low-precision inertial navigation system relative to the North-East coordinate system based on the pitch angle, roll angle, and yaw angle of the high- and low-precision inertial navigation systems: Based on the attitude matrices of the high-precision and low-precision inertial navigation systems relative to the North-East coordinate system, the relative installation error matrices of the high-precision inertial navigation system X1Y1Z1 and the low-precision inertial navigation system X2Y2Z2 are calculated: 。 3. The method for calibrating and compensating relative installation errors of a multi-pose measuring device according to claim 2, characterized in that, Step S4 includes: Multi-source integrated navigation systems use a high-precision inertial navigation body coordinate system as the coordinate reference; Through the relative installation error matrix Compensation is used to adjust the attitude matrix measured dynamically in real time by the low-precision inertial navigation system. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system: In the formula: To compensate for low-precision inertial navigation real-time dynamic measurement attitude matrix after installation error; From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion. Through the relative installation error matrix Compensation is applied to the attitude matrix measured dynamically in real time by the star sensor. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system: In the formula: The attitude matrix of the star sensor is dynamically measured in real time to compensate for installation errors; From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

4. A system for calibrating and compensating relative installation errors of a multi-pose measuring device, characterized in that, include: Module M1: Measures the star sensor attitude and transforms coordinates; Module M2: Calibrates the relative installation error of high-precision inertial navigation and star sensors; Module M3: Calibrates the relative installation error of high-precision and low-precision inertial navigation systems for self-alignment; Module M4: Compensates for relative installation errors; The module M1 includes: With the carrier platform's X-axis pointing vertically upwards, the star sensor performs attitude measurements at night with its orientation also pointing vertically upwards, obtaining the attitude quaternions from the star sensor's intrinsic frame to the J2000 inertial frame. ; Based on the obtained attitude transformation from the star-sensitive datum to the J2000 inertial frame, the attitude array of the star-sensitive datum to the North-East coordinate system is converted. The specific conversion process is as follows: First, from the star-aware system to the attitude quaternion of the J2000 inertial frame. The attitude array is constructed as follows: In the formula, , , , These are the four elements of the attitude quaternion from the star-aware system to the J2000 inertial frame; Secondly, the transformation matrix from the J2000 inertial frame to the Earth coordinate system is: In the formula, A The transformation matrix between the quasi-Earth-fixed coordinate system and the Earth coordinate system. B This is the Earth's rotation transformation matrix. C Nucleation transformation matrix, D This is the precession transformation matrix; Finally, the process of converting the attitude array from the star-sensitive intrinsic system to the J2000 inertial frame to the attitude array from the star-sensitive intrinsic system to the North-East coordinate system is as follows: in, In the formula, Earth-to-navigation-system conversion matrix, For the longitude of the carrier, L The latitude of the carrier; The module M2 includes: The carrier platform's X-axis is vertically upward, as are the star sensor's X3 axis and the high-precision inertial navigation's X1 axis. The star sensor completes attitude measurement and acquires the star sensor attitude array from the local system to the north-sky-east coordinate system according to step S1. ; The high-precision inertial navigation system (INS) performs self-alignment based on the bound geographic parameters. After alignment, the attitude quaternions of the INS relative to the North-Sky-East coordinate system are obtained, and the attitude matrix is ​​obtained accordingly. In the formula, , , , These are the four elements of the attitude quaternion of the high-precision inertial navigation system relative to the North-East coordinate system; Obtain the relative installation error matrix between the high-precision inertial navigation system X1Y1Z1 and the star sensor system X3Y3Z3: 。 5. The multi-pose measurement equipment relative installation error calibration and compensation system according to claim 4, characterized in that, The module M3 includes: a carrier platform placed horizontally, with the high-precision inertial navigation Y1 axis and the low-precision inertial navigation Y2 axis both pointing vertically upwards, respectively self-aligned to obtain the high-precision inertial navigation pitch angle. Roll angle and low-precision inertial navigation pitch angle Roll angle ; The carrier platform is rotated 90° and placed horizontally, with the high-precision inertial navigation Z1 axis and the low-precision inertial navigation Z2 axis both pointing vertically upwards, and self-aligned to obtain the high-precision inertial navigation pitch angle. and low-precision inertial navigation pitch angle ; When the carrier platform is rotated 90° and placed horizontally, the high-precision and low-precision inertial navigation pitch angles are the same as the high-precision and low-precision inertial navigation yaw angles when the carrier platform is normally placed horizontally, that is: Calculate the attitude matrix of the high- and low-precision inertial navigation system relative to the North-East coordinate system based on the pitch angle, roll angle, and yaw angle of the high- and low-precision inertial navigation systems: Based on the attitude matrices of the high-precision and low-precision inertial navigation systems relative to the North-East coordinate system, the relative installation error matrices of the high-precision inertial navigation system X1Y1Z1 and the low-precision inertial navigation system X2Y2Z2 are calculated: 。 6. The multi-pose measurement equipment relative installation error calibration and compensation system according to claim 5, characterized in that, The module M4 includes: Multi-source integrated navigation systems use a high-precision inertial navigation body coordinate system as the coordinate reference; Through the relative installation error matrix Compensation is used to adjust the attitude matrix measured dynamically in real time by the low-precision inertial navigation system. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system: In the formula: To compensate for low-precision inertial navigation real-time dynamic measurement attitude matrix after installation error; From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion. Through the relative installation error matrix Compensation is applied to the attitude matrix measured dynamically in real time by the star sensor. Attitude matrix projected onto the high-precision inertial navigation body coordinate system relative to the North-East coordinate system: In the formula: The attitude matrix of the star sensor is dynamically measured in real time to compensate for installation errors; From attitude array After compensation, pitch angle, roll angle, and yaw angle are extracted and used for multi-source integrated navigation information fusion.

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