Aircraft attitude continuous correction method and system based on star navigation device

Abstract

The present application discloses a method and system for continuous correction of an aircraft's attitude based on a starlight navigation device, which relates to the field of test and measurement. The method comprises: using a starlight navigation device to perform a stargazing process as an iterative cycle, acquiring aircraft attitude observation data within each iterative cycle; and using a linear Kalman filter to perform online estimation and correction of all error terms in the attitude chain based on the aircraft's attitude observation data. Correction of all error terms in the attitude chain is completed until the corrected error terms are less than a set threshold. This application can improve the accuracy of aircraft attitude correction.

Description

Technical Field

[0001] The present application relates to the field of testing and measurement, and in particular to a method and system for continuous correction of an aircraft attitude based on a starlight navigation device. Background Art

[0002] As a branch of inertial navigation system research, attitude estimation, whose parameters are reliable, is crucial for ensuring the safe and efficient operation of moving vehicles such as aircraft. Inertial navigation systems use gyroscopes and accelerometers to measure the vehicle's angular velocity and linear velocity. Through navigation solutions, they continuously output complete navigation parameter information, including position, velocity, and attitude. However, over long periods of continuous operation, the errors of pure inertial navigation systems accumulate and gradually diverge over time. Starlight navigation devices can determine the vehicle's attitude information within the inertial system by observing star vectors, offering the advantages of high accuracy and zero error accumulation. Starlight navigation devices combine inertial navigation with astronomical navigation, utilizing the high-precision attitude information provided by star sensors to correct inertial navigation system errors, thereby achieving high-precision navigation.

[0003] Nowadays, starlight navigation systems are increasingly adopting a fully strapdown mode, meaning the strapdown inertial navigation system and star sensors are installed in a strapdown manner. This system's error sources primarily fall into three categories: platform error angle, introduced during the initial alignment process and remaining constant during navigation; gyroscopic error, primarily consisting of gyro zero error, scale error, and installation error, which accumulates over time during navigation and introduces platform error angle; and starlight installation error, caused by misalignment between the star sensor and the inertial navigation system. These three error sources are coupled and transmitted during the integration process. Inertial / starlight integration technology primarily uses star sensor measurement information to directly correct the inertial navigation system, lacking the ability to model and decouple all system error terms, resulting in reduced integration accuracy. To meet the requirements of high-precision aircraft attitude determination, continuous attitude correction based on decoupled identification of all error terms is crucial, but no relevant solutions have been proposed. Summary of the Invention

[0004] The purpose of this application is to provide a method and system for continuous correction of aircraft attitude based on a starlight navigation device, which can improve the correction accuracy of aircraft attitude.

[0005] To achieve the above objectives, this application provides the following solutions:

[0006] In a first aspect, the present application provides a method for continuous correction of an aircraft attitude based on a starlight navigation device, comprising:

[0007] During each iteration, the aircraft's attitude observation data is acquired; the aircraft's attitude observation data includes: an inertial navigation system installation reference vector, a starlight navigation system installation reference vector, and an attitude matrix of the inertial navigation system relative to the navigation system. One stargazing process using the starlight navigation system constitutes one iteration.

[0008] A linear Kalman filter is used to complete online estimation and correction of the attitude chain full error term based on the attitude observation data of the aircraft, and the attitude chain full error term correction is completed when the corrected attitude chain full error term is less than a set threshold.

[0009] Optionally, the attitude chain full error term includes: gyro error, platform error angle and satellite inertial installation error.

[0010] Optionally, a starlight navigation device is used to perform the stargazing process based on a static or quasi-static stargazing strategy.

[0011] Optionally, the measurement equation of the linear Kalman filter is expressed as:

[0012] ;

[0013] Where, is the platform error angle of the k+1th iteration cycle, is the satellite installation error of the k+1th iteration period, is the inertial navigation installation reference vector of the k+1th iteration cycle, is the installation reference vector of the star navigation device in the k+1th iteration cycle, is the identity matrix, is the attitude matrix of the inertial navigation system relative to the navigation system, is the right multiplication symbol.

[0014] Optionally, the maximum value of the set threshold is 5 arc seconds.

[0015] Optionally, the attitude of the aircraft is corrected using the corrected attitude chain full error term, including:

[0016] Determine the components of the corrected inertial installation error in the three axes of the navigation system;

[0017] The attitude angle of the aircraft is corrected based on the three axial components of the corrected satellite inertial installation error in the navigation system and the corrected platform error angle to complete the attitude correction of the aircraft; the attitude angle includes: pitch angle, roll angle and heading angle.

[0018] Optionally, the number of iteration cycles is at least 7.

[0019] In a second aspect, the present application provides an aircraft attitude continuous correction system based on a starlight navigation device, comprising: a starlight navigation device, a dual-axis indexing mechanism, a navigation computer, an optical theodolite, a first fixed light source, and a second fixed light source;

[0020] The starlight navigation device exchanges data with the navigation computer; the starlight navigation device is mounted on the dual-axis rotation mechanism; the dual-axis rotation mechanism is mounted on the body of the aircraft; the first fixed light source and the second fixed light source are arranged on the ground at a set angle;

[0021] The optical theodolite is used to provide a correction reference; the first fixed light source and the second fixed light source are both used to simulate starlight; the starlight navigation device, driven by the dual-axis indexing mechanism, combines the first fixed light source, the second fixed light source and the correction reference to perform a stargazing process; the navigation computer is used to implement the above-mentioned method for continuous correction of aircraft attitude based on the starlight navigation device to achieve continuous correction of aircraft attitude.

[0022] Optionally, the star navigation device includes a star sensor.

[0023] Optionally, the aircraft attitude continuous correction system based on the starlight navigation device further comprises: a communication interface;

[0024] The starlight navigation device exchanges data with the navigation computer via the communication interface.

[0025] According to the specific embodiments provided in this application, this application has the following technical effects:

[0026] The present application provides a method and system for continuous correction of an aircraft's attitude based on a starlight navigation device. By using a linear Kalman filter to complete online estimation and correction of the full error term of the attitude chain based on the aircraft's attitude observation data, the full error term of the system can be modeled and decoupled for identification, thereby resolving the problem of degradation of the combined accuracy. Furthermore, within each iteration cycle, a linear Kalman filter is used to complete online estimation and correction of the full error term of the attitude chain based on the aircraft's attitude observation data until the corrected full error term of the attitude chain is less than a set threshold. At this point, correction of the full error term of the attitude chain is completed, thereby achieving continuous correction of the aircraft's attitude, improving the correction accuracy of the aircraft's attitude, and meeting the requirements of high-precision attitude determination. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0028] Figure 1 A flowchart of a method for continuous correction of an aircraft attitude based on a starlight navigation device provided in one embodiment of the present application;

[0029] Figure 2 A schematic structural diagram of a system for continuous correction of an aircraft attitude based on a starlight navigation device according to an embodiment of the present application.

[0030] Figure 3 A schematic diagram of a light source angle measurement process according to an embodiment of the present application;

[0031] Figure 4 A schematic diagram of the stargazing process provided in one embodiment of the present application.

[0032] Reference numerals:

[0033] 11-External frame, 12-Inner frame, 13-Star navigation device, 14-Communication interface, 15-Dual-axis rotation mechanism, 16-Navigation computer, 17-Star sensor, 18-Inner platform. DETAILED DESCRIPTION

[0034] The following will be combined with the drawings in the embodiments of this application to clearly and completely describe the technical solutions in the embodiments of this application. Obviously, the embodiments described are only part of the embodiments of this application, not all of the embodiments. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of this application.

[0035] In order to make the above-mentioned purposes, features and advantages of the present application more obvious and easy to understand, the present application is further described in detail below with reference to the accompanying drawings and specific implementation methods.

[0036] In an exemplary embodiment, the present application provides a method for continuous correction of aircraft attitude based on a starlight navigation device. The method is executed by a computer device, specifically a computer device such as a terminal or a server, or a terminal and a server. In the embodiment of the present application, the method is applied to a server as an example for explanation. Figure 1 As shown, the method includes:

[0037] Step 100: Acquire aircraft attitude observation data during each iteration. This aircraft attitude observation data includes the inertial navigation system installation reference vector, the star navigation system installation reference vector, and the attitude matrix of the inertial navigation system relative to the navigation system. One star observation process using the star navigation system constitutes one iteration. For example, the star navigation system can be used to perform star observations using a static or quasi-static strategy.

[0038] Step 101: Use a linear Kalman filter to perform online estimation and correction of all attitude chain error terms based on the aircraft's attitude observation data. Correction is completed when the corrected attitude chain error terms are less than a set threshold (e.g., 5 arc seconds). The attitude chain error terms include 15 items, including gyro error, platform error angle, and satellite inertial error.

[0039] In another exemplary embodiment of the present application, based on the above-mentioned method for continuous correction of aircraft attitude, it is assumed that 15 errors such as gyro error, platform error angle and satellite inertial installation error can be observed through 7 star observation processes (i.e., 7 iteration cycles), thereby establishing a full attitude link error state model. The linear Kalman filter is used as the optimal estimation method to complete the online estimation of all error terms in the attitude chain and the correction of attitude angle. Based on this, the platform error angle is used as the optimal estimation method. Installation error of star habit Taking the specific iterative correction process as an example, the implementation process of the above steps 100 and 101 provided in this application is explained.

[0040] 1) The first stargazing process (i.e. the first observation).

[0041] The star vector observation equation considering the inertial installation error is:

[0042] (1)

[0043] Where, is the inertial navigation installation reference vector of the first iteration cycle, is the installation reference vector of the star navigation device in the first iteration cycle, is the identity matrix, is the platform error angle of the first iteration cycle, is the attitude matrix of the inertial navigation system relative to the navigation system, It is the satellite inertial installation error defined in the inertial navigation system of the first iteration cycle.

[0044] Calculate the platform error angle using a linear Kalman filter Installation error of star habit When , the measurement equation of the linear Kalman filter is expressed as:

[0045] (2)

[0046] Where, is the right multiplication symbol.

[0047] The observation results obtained through the first stargazing process can be used to obtain the platform error angle of the first iteration cycle. Installation error of star habit , and use this to correct the attitude angle solved in real time, we have:

[0048] (3)

[0049] Where, and are the corrected value and original value of the pitch angle in the first iteration cycle, and are the corrected value and original value of the roll angle in the first iteration cycle, and are the corrected value and original value of the heading angle in the first iteration cycle, 、 and They are the platform error angle vectors of the first iteration cycle In the navigation system The three axial components, 、 and They are the star installation errors of the first iteration cycle In the navigation system The three axial components compensate for the platform error angle vector in the navigation system The three axial components and the satellite inertial installation error are in the navigation system Three axial components. Then determine the platform error angle Installation error of star habit Are they all less than 5 arc seconds? If so, the iteration is completed; otherwise, the next iteration cycle is performed.

[0050] 2) The second stargazing process (i.e. the second observation).

[0051] The star vector observation equation considering the inertial installation error is:

[0052] (4)

[0053] Where, is the inertial navigation installation reference vector of the second iteration cycle, is the installation reference vector of the star navigation device in the second iteration cycle, is the platform error angle of the second iteration cycle, It is the satellite inertial installation error defined in the inertial navigation system of the second iteration cycle.

[0054] Next, the platform error angle is calculated using a linear Kalman filter. Installation error of star habit ,have:

[0055] (5)

[0056] The platform error angle of the second iteration cycle can be obtained through the observation results of the second stargazing process. Installation error of star habit , and use this to correct the attitude angle solved in real time, we have:

[0057] (6)

[0058] Where, and are the corrected value and original value of the pitch angle in the second iteration cycle, and are the corrected value and original value of the roll angle in the second iteration cycle, and are the corrected value and original value of the heading angle in the second iteration cycle, 、 and They are the platform error angle vectors of the second iteration cycle In the navigation system The three axial components, 、 and They are the second iteration cycle star inertia installation errors In the navigation system Three axial components.

[0059] Determine the platform error angle Installation error of star habit Are they all less than 5 arc seconds? If so, the iteration is completed; otherwise, the next iteration cycle is performed.

[0060] 3) The third stargazing process (i.e. the third observation).

[0061] The star vector observation equation considering the inertial installation error is:

[0062] (7)

[0063] Where, is the inertial navigation installation reference vector of the third iteration cycle, is the installation reference vector of the star navigation device in the third iteration cycle, is the platform error angle of the third iteration cycle, It is the satellite inertial installation error defined in the inertial navigation system of the third iteration cycle.

[0064] Next, the platform error angle is calculated using a linear Kalman filter. Installation error of star habit ,have:

[0065] (8)

[0066] The platform error angle of the third iteration cycle can be obtained through the third observation result. Installation error of star habit , and use this to correct the attitude angle solved in real time, we have:

[0067] (9)

[0068] Where, and are the corrected value and original value of the pitch angle in the third iteration cycle, and are the corrected value and original value of the roll angle in the third iteration cycle, and are the corrected value and original value of the heading angle in the third iteration cycle, 、 and They are the platform error angle vectors in the third iteration cycle In the navigation system The three axial components, 、 and The third iteration cycle star inertia installation error is In the navigation system Three axial components.

[0069] Then determine the platform error angle Installation error of star habit Are they all less than 5 arc seconds? If so, the iteration is completed; otherwise, the next iteration cycle is performed.

[0070] And so on, until the k+1th stargazing process (i.e. the k+1th observation), we have:

[0071] The star vector observation equation considering the satellite inertial installation error is shown in the following formula (10).

[0072] (10)

[0073] Where, is the inertial navigation installation reference vector of the k+1th iteration cycle, is the installation reference vector of the star navigation device in the k+1th iteration cycle, is the platform error angle of the k+1th iteration cycle, It is the satellite inertial installation error defined in the inertial navigation system of the k+1th iteration cycle.

[0074] Next, the platform error angle is calculated using a linear Kalman filter. Installation error of star habit ,have:

[0075] (11)

[0076] The platform error angle of the k+1th iteration cycle can be obtained through the observation results of the k+1th stargazing process. Installation error of star habit , and use this to correct the attitude angle solved in real time, we have:

[0077] (12)

[0078] Where, and are the corrected value and original value of the pitch angle in the k+1th iteration period, and are the corrected value and original value of the roll angle in the k+1th iteration period, and are the corrected value and original value of the heading angle in the k+1th iteration period, 、 and They are the platform error angle vectors of the k+1th iteration period In the navigation system The three axial components, 、 and They are the star installation errors in the k+1th iteration cycle In the navigation system Three axial components.

[0079] Then determine the platform error angle Installation error of star habit Are they all less than 5 arc seconds? If so, the iteration is completed; otherwise, the next iteration cycle is performed.

[0080] Based on the above description, in step 101, the process of correcting the attitude of the aircraft using the corrected attitude chain full error term can be described as follows:

[0081] 1. Determine the components of the corrected inertial installation error in the three axes of the navigation system.

[0082] 2. Based on the corrected satellite inertial installation error in the three axes of the navigation system and the corrected platform error angle, the aircraft's attitude angle is corrected to complete the aircraft's attitude correction. The attitude angles include pitch, roll, and heading.

[0083] In summary, this application further improves the attitude accuracy of the aircraft by correcting the full error terms of the attitude chain of the aircraft system and continuously correcting the aircraft attitude. It can solve the problems of deterioration of combination accuracy and low attitude accuracy of the aircraft system caused by the aircraft system's lack of ability to model and decouple and identify the full error terms of the system.

[0084] Based on the same inventive concept, embodiments of the present application also provide a system for continuously correcting an aircraft's attitude based on a starlight navigation device, for implementing the aforementioned method for continuously correcting an aircraft's attitude. The solution provided by this system is similar to the solution described in the aforementioned method. Therefore, the specific limitations of one or more embodiments of the system for continuously correcting an aircraft's attitude based on a starlight navigation device provided below can be found in the aforementioned definition of the method for continuously correcting an aircraft's attitude, and will not be further elaborated here.

[0085] In an exemplary embodiment, Figure 2 As shown, a system for continuous correction of aircraft attitude based on a starlight navigation device is provided, comprising: a starlight navigation device 13, a dual-axis indexing mechanism 15, a navigation computer 16, an optical theodolite (not shown), a first fixed light source (i.e. Figure 2 Light source 1 in the image) and a second fixed light source (i.e. Figure 2 The star navigation device 13 includes a star sensor 17.

[0086] The starlight navigation device 13 exchanges data with the navigation computer 16. The starlight navigation device 13 is mounted on a dual-axis rotation mechanism. The dual-axis rotation mechanism 15 is mounted on the aircraft body. The first fixed light source and the second fixed light source are set on the ground at a set angle.

[0087] An optical theodolite is used to provide a correction reference. For example, the optical theodolite is used to measure the azimuth of the North Star to obtain the true north direction, which is then aligned on the surface of a prism on the pier, namely the north reference prism. After alignment, the north reference prism is fixed to the pier to serve as the north reference (i.e., the correction reference). Both the first and second fixed light sources are used to simulate starlight. Driven by the dual-axis indexing mechanism 15, the starlight navigation device 13 combines the first and second fixed light sources with the correction reference to perform stargazing. The navigation computer 16 is used to implement the aforementioned method for continuous aircraft attitude correction based on the starlight navigation device to achieve continuous correction of the aircraft's attitude.

[0088] As an optional implementation, Figure 2 As shown, the starlight navigation device 13 exchanges data with the navigation computer 16 via the communication interface 14. The communication interface 14 may be a 1553b communication interface.

[0089] As an optional implementation, during the actual installation process, if Figure 2 As shown, the aircraft attitude continuous correction system provided by the present application can be installed inside the nose of an aircraft. The dual-axis indexing mechanism 15 rotates under the drive of the inner frame 12 and the outer frame 11. The inner platform 18 is a carrier for installing the starlight navigation device 13.

[0090] In an exemplary embodiment, the continuous attitude correction system for aircraft is used to implement continuous attitude correction. It is necessary to first decouple the full error term of the attitude chain before performing continuous attitude correction. Therefore, in this embodiment, the continuous attitude correction process includes:

[0091] Step 1: Install the system: Install the starlight navigation device 13 on the dual-axis rotation mechanism 15. The dual-axis rotation mechanism 15 can realize the attitude adjustment of the starlight navigation device 13. Install two fixed light sources (i.e., the first fixed light source and the second fixed light source) on a stable base on the ground to simulate real starlight and generate Figure 2 The star vector 1 and star vector 2 are shown. In order to make the error amplification factor close to 1, the angle between the two fixed light sources It can be set to 90°. In order to ensure that the parallel light beam of the fixed light source can completely enter the field of view of the star sensor 17, it is necessary to ensure that the center height of the fixed light source is basically consistent with the center height of the field of view of the star sensor 17 installed on the dual-axis rotation mechanism 15, and the simulated star vector needs to be calibrated and measured before the test. The navigation computer 16 is connected to the star navigation device 13 through the 1553b communication interface to complete data synchronization and real-time acquisition, and is used for navigation solution. The azimuth of the North Star is measured using an optical theodolite to obtain the true north direction, and it is directed to the prism on the pier, that is, the north reference prism, and collimated. After collimation, the north reference prism is fixed on the pier to serve as the north reference. The star sensor 17 corrects the attitude error and device error of the star navigation device 13 by measuring known simulated starlight information.

[0092] Step 2: Establish the attitude reference of the fixed light source: Use the optical theodolite and the north reference to directly measure the first fixed light source (i.e. Figure 2 By adjusting the attitude angle of the indexing mechanism, the parallel beam of the first fixed light source falls exactly near the center point of the imaging plane of the star sensor 17, ensuring high-precision measurement of the starlight vector. Then, the outer frame 11 of the dual-axis rotation mechanism 15 is rotated so that the second fixed light source (i.e. Figure 2 The parallel beam of light source 2) falls exactly on the last imaging point in the imaging plane of star sensor 17. The rotation angle needs to be calculated by reading the output of star sensor 17 in real time. When star sensor 17 points to the first fixed light source, the recorded azimuth angle of the indexing mechanism is , when pointing to the second fixed light source, the azimuth angle of the indexing mechanism is recorded as , so the angle between two fixed light sources can be expressed as , The azimuth angle of the second fixed light source is obtained by measuring the azimuth angle of the first fixed light source and the light source angle. The light source angle measurement process is as follows: Figure 3 shown. Figure 3 In , FP stands for focal plane.

[0093] Step 3: Analysis of stargazing times: For the star navigation device 13, the attitude chain error mainly includes 15 errors, including gyro zero position, gyro scale, gyro installation error, star inertial installation error, platform error angle, etc. For the small field of view star sensor 17, the star inertial installation error around the optical axis is less observable. Therefore, at least 7 stargazing times are required to make all error terms observable.

[0094] Step 4, stargazing process: adopt static or quasi-static stargazing strategy, that is, after the attitude adjustment is completed, the starlight navigation device 13 is in a completely static condition to complete the stargazing. The starlight navigation device 13 needs to be initially aligned for 5 minutes, and then the stargazing is carried out 7 times in sequence as the indexing mechanism rotates. The specific stargazing process is as follows: Figure 4 shown. Figure 4 In the embodiment, the time for measuring light source 1 and light source 2, and the values ​​of the rotation angles of the inner frame and the outer frame are not specifically limited to this application and are determined according to measurement requirements during actual application.

[0095] Step 5: Correct all attitude chain errors: The seven stargazing processes described above allow observation of 15 errors, including gyro errors, platform error angles, and satellite inertial error, to establish a full attitude chain error state model. A linear Kalman filter is used as the optimal estimation method to perform online estimation and correction of all attitude chain error terms.

[0096] Step 6: Continuous attitude correction: Each observation is an iteration cycle, and the measurement equation of the linear Kalman filter shown in formula (2), formula (5), formula (8) or formula (11) is repeatedly executed to obtain the platform error angle Installation error of star habit And compensate it, and correct the attitude angle calculated in real time. The iterative process continues until the platform error angle Installation error of star habit When the deviations are less than 5 arc seconds, the aircraft stops, thus realizing continuous correction of its attitude.

[0097] In an exemplary embodiment, a computer device is provided, which may be a server or a terminal. The computer device includes a processor, a memory, an input / output (I / O) interface, and a communication interface. The processor, memory, and I / O interface are connected via a system bus, and the communication interface is connected to the system bus via the I / O interface. The processor of the computer device is configured to provide computing and control capabilities. The memory of the computer device includes a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system, a computer program, and a database. The internal memory provides an environment for the operation of the operating system and computer program in the non-volatile storage medium. The database of the computer device is configured to store continuous aircraft attitude correction data based on a starlight navigation device. The I / O interface of the computer device is configured to exchange information between the processor and an external device. The communication interface of the computer device is configured to communicate with an external terminal via a network connection. When executed by the processor, the computer program implements a method for continuous aircraft attitude correction based on a starlight navigation device.

[0098] In an exemplary embodiment, a computer device is provided, including a memory and a processor. The memory stores a computer program, and the processor implements the steps in the above method embodiments when executing the computer program.

[0099] In an exemplary embodiment, a computer-readable storage medium is provided, storing a computer program. When the computer program is executed by a processor, the steps in the above-mentioned method embodiments are implemented.

[0100] In an exemplary embodiment, a computer program product is provided, including a computer program. When the computer program is executed by a processor, the steps in the above method embodiments are implemented.

[0101] It should be noted that the user information (including but not limited to user device information, user personal information, etc.) and data (including but not limited to data used for analysis, stored data, displayed data, etc.) involved in this application are all information and data authorized by the user or fully authorized by all parties, and the collection, use and processing of relevant data must comply with relevant regulations.

[0102] Those skilled in the art will appreciate that all or part of the processes in the above-mentioned embodiments can be implemented by instructing the relevant hardware through a computer program. The computer program can be stored in a non-volatile computer-readable storage medium. When the computer program is executed, it can include the processes of the above-mentioned embodiments. In particular, any reference to memory, database, or other media used in the embodiments provided in this application can include at least one of non-volatile and volatile memory. Non-volatile memory can include read-only memory (ROM), magnetic tape, floppy disk, flash memory, optical memory, high-density embedded non-volatile memory, resistive random access memory (ReRAM), magnetic random access memory (MRAM), ferroelectric random access memory (FRAM), phase change memory (PCM), graphene memory, etc. Volatile memory can include random access memory (RAM) or external cache memory, etc. By way of illustration and not limitation, RAM may be in various forms, such as static random access memory (SRAM) or dynamic random access memory (DRAM).

[0103] The databases involved in the various embodiments provided herein may include at least one of a relational database and a non-relational database. Non-relational databases may include, but are not limited to, distributed databases based on blockchains. The processors involved in the various embodiments provided herein may include, but are not limited to, general-purpose processors, central processing units, graphics processing units, digital signal processors, programmable logic units, data processing logic units based on quantum computing, and the like.

[0104] The technical features of the above embodiments can be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0105] This document uses specific examples to illustrate the principles and implementation methods of this application. The description of the above examples is only intended to help understand the method and core concept of this application. At the same time, for those skilled in the art, based on the concept of this application, there may be changes in the specific implementation methods and application scope. In summary, the content of this specification should not be understood as limiting this application.

Claims

1. A method for continuous correction of aircraft attitude based on a star navigation device, characterized in that: include: In each iteration cycle, the attitude observation data of the aircraft is obtained; The attitude observation data of the aircraft includes: an inertial navigation installation reference vector, a starlight navigation device installation reference vector, and an attitude matrix of the inertial navigation system relative to the navigation system; a starlight navigation device is used to perform a star observation process as an iteration cycle; A linear Kalman filter is used to complete the online estimation and correction of the attitude chain full error term based on the attitude observation data of the aircraft, until the corrected attitude chain full error term is less than the set threshold, and the attitude chain full error term correction is completed; the measurement equation of the linear Kalman filter is expressed as: ； Where, is the platform error angle of the k+1th iteration cycle, is the satellite installation error of the k+1th iteration period, is the inertial navigation installation reference vector of the k+1th iteration cycle, is the installation reference vector of the star navigation device in the k+1th iteration cycle, is the identity matrix, is the attitude matrix of the inertial navigation system relative to the navigation system, is the right multiplication symbol.

2. The method for continuous correction of aircraft attitude based on a starlight navigation device according to claim 1, characterized in that: The attitude chain full error terms include: gyro error, platform error angle and satellite inertial installation error.

3. The method for continuous correction of aircraft attitude based on a star navigation device according to claim 1, characterized in that: The starlight navigation device is used to conduct stargazing based on a static or quasi-static stargazing strategy.

4. The method for continuous correction of aircraft attitude based on a starlight navigation device according to claim 1, characterized in that: The maximum value of the set threshold is 5 arc seconds.

5. The method for continuous correction of aircraft attitude based on a starlight navigation device according to claim 2, characterized in that: The attitude of the aircraft is corrected using the corrected attitude chain full error terms, including: Determine the components of the corrected inertial installation error in the three axes of the navigation system; The attitude angle of the aircraft is corrected based on the three axial components of the corrected satellite inertial installation error in the navigation system and the corrected platform error angle to complete the attitude correction of the aircraft; the attitude angle includes: pitch angle, roll angle and heading angle.

6. The method for continuous correction of aircraft attitude based on a starlight navigation device according to claim 1, characterized in that: The number of the iteration cycles is at least 7.

7. A continuous aircraft attitude correction system based on a starlight navigation device, characterized in that: The aircraft attitude continuous correction system based on the starlight navigation device comprises: a starlight navigation device, a dual-axis indexing mechanism, a navigation computer, an optical theodolite, a first fixed light source and a second fixed light source; The starlight navigation device exchanges data with the navigation computer; the starlight navigation device is mounted on the dual-axis rotation mechanism; the dual-axis rotation mechanism is mounted on the body of the aircraft; the first fixed light source and the second fixed light source are arranged on the ground at a set angle; The optical theodolite is used to provide a correction reference; the first fixed light source and the second fixed light source are both used to simulate starlight; the starlight navigation device, driven by the dual-axis rotation mechanism, combines the first fixed light source, the second fixed light source and the correction reference to perform a stargazing process; the navigation computer is used to implement the continuous correction method for aircraft attitude based on the starlight navigation device as described in any one of claims 1 to 6 to achieve continuous correction of the aircraft attitude.

8. The aircraft attitude continuous correction system based on the starlight navigation device according to claim 7, characterized in that: The starlight navigation device includes a star sensor.

9. The aircraft attitude continuous correction system based on the star navigation device according to claim 7, characterized in that: The aircraft attitude continuous correction system based on the starlight navigation device also includes: a communication interface; The starlight navigation device exchanges data with the navigation computer via the communication interface.

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