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

Abstract

The invention discloses an aircraft attitude continuous correction method and system based on a starlight navigation device, and relates to the field of test and measurement, and the method comprises the steps: employing the starlight navigation device to carry out a star observation process as an iteration period, and obtaining the attitude observation data of an aircraft in each iteration period; and adopting a linear Kalman filter to complete online estimation and correction of the attitude chain total error term based on the attitude observation data of the aircraft until the corrected attitude chain total error term is smaller than a set threshold value, and completing correction of the attitude chain total error term. The correction precision of the aircraft attitude can be improved.

Description

Technical Field

[0001] This application relates to the field of test and measurement, and particularly to a method and system for continuously correcting the attitude of an aircraft based on a starlight navigation device. Background Art

[0002] Attitude estimation, as a branch field of studying inertial navigation systems, the reliability of its parameters is the core to ensure the safe and efficient operation of moving carriers such as aircraft. The inertial navigation system uses gyroscopes and accelerometers to measure the rotational angular velocity and linear velocity of the carrier, and can continuously output complete navigation parameter information through navigation calculation, including position, velocity, and attitude. However, during long-term continuous operation, the errors of a pure inertial navigation system will gradually diverge over time. The starlight navigation device can determine the attitude information of the carrier in the inertial system by observing the star vectors, and has the advantages of high precision and non-accumulating errors. The starlight navigation device combines inertial navigation and astronomical navigation, and uses the high-precision attitude information provided by the star sensor to correct the errors of the inertial navigation system, thereby achieving high-precision navigation.

[0003] Nowadays, starlight navigation devices are gradually trending towards a fully strapdown mode, that is, the strapdown inertial navigation system and the star sensor are installed in a strapdown manner. The error sources of this system mainly include three categories. One is the platform error angle, which is introduced during the initial alignment process and remains unchanged during navigation. The second is the gyro error, which mainly includes the gyro zero position error, scale error, and installation error, and introduces the platform error angle during navigation as time accumulates. The third is the star-inertial installation error, which is caused by the installation deviation between the star sensor and the inertial navigation system. These three error sources are coupled and transmitted during the combination process, and the inertial / starlight combination technology mainly directly corrects the inertial navigation system with the measurement information of the star sensor, and does not have the ability to model and decouple and identify all error terms of the system, resulting in the deterioration of the combination accuracy. In view of the high-precision attitude determination requirements of aircraft, it is of great significance to carry out research on continuous attitude correction based on the decoupling and identification of all error terms, but currently no solutions related to this research have been proposed. Summary of the Invention

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

[0005] To achieve the above purpose, this application provides the following solutions: In the first aspect, this application provides a method for continuously correcting the attitude of an aircraft based on a starlight navigation device, including: In each iteration period, obtain the attitude observation data of the aircraft; the attitude observation data of the aircraft includes: the inertial navigation installation reference vector, the starlight navigation device installation reference vector, and the attitude matrix of the inertial navigation body system relative to the navigation system; one star observation process using the starlight navigation device is one iteration period; Use a linear Kalman filter to complete the online estimation and correction of the full error terms of the attitude chain based on the attitude observation data of the aircraft. When the full error terms of the corrected attitude chain are less than the set threshold, complete the correction of the full error terms of the attitude chain.

[0006] Optionally, the full error terms of the attitude chain include: gyro error, platform error angle, and star-inertial installation error.

[0007] Optionally, use the starlight navigation device to perform the star observation process based on a static or quasi-static star observation strategy.

[0008] Optionally, the measurement equation of the linear Kalman filter is expressed as: ; In the formula, is the platform error angle in the (k + 1)-th iteration period, is the star-inertial installation error in the (k + 1)-th iteration period, is the inertial navigation installation reference vector in the (k + 1)-th iteration period, is the starlight navigation device installation reference vector in the (k + 1)-th iteration period, is the identity matrix, is the attitude matrix of the inertial navigation body system relative to the navigation system, is the right multiplication symbol.

[0009] Optionally, the maximum value of the set threshold is 5 arcseconds.

[0010] Optionally, use the corrected full error terms of the attitude chain to correct the attitude of the aircraft, including: Determine the components of the corrected star-inertial installation error in the three axes of the navigation system; Based on the components of the corrected star-inertial installation error in the three axes of the navigation system and the corrected platform error angle, correct the attitude angles of the aircraft to complete the attitude correction of the aircraft; the attitude angles include: pitch angle, roll angle, and heading angle.

[0011] Optionally, the number of iteration periods is at least 7.

[0012] In a second aspect, the present application provides an aircraft attitude continuous correction system based on a starlight navigation device, including: a starlight navigation device, a two-axis indexing mechanism, a navigation computer, an optical theodolite, a first fixed light source, and a second fixed light source; The starlight navigation device performs data interaction with the navigation computer; the starlight navigation device is installed on the two-axis indexing mechanism; the two-axis indexing mechanism is installed on the airframe 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; both the first fixed light source and the second fixed light source are used to simulate starlight; the starlight navigation device, driven by the two-axis indexing mechanism, combines the first fixed light source, the second fixed light source and the correction reference to perform the star observation process; the navigation computer is used to implement the aircraft attitude continuous correction method provided above based on the starlight navigation device to achieve continuous correction of the aircraft attitude.

[0013] Optionally, the starlight navigation device includes a star sensor.

[0014] Optionally, the aircraft attitude continuous correction system based on the starlight navigation device further includes: a communication interface; The starlight navigation device performs data interaction with the navigation computer through the communication interface.

[0015] According to the specific embodiments provided in this application, this application has the following technical effects: This application provides an aircraft attitude continuous correction method and system based on a starlight navigation device. By using a linear Kalman filter to complete the online estimation and correction of all error terms of the attitude chain based on the attitude observation data of the aircraft, the system can model and decouple and identify all error terms of the system, thereby solving the problem of deterioration of the combined accuracy. And, in each iteration cycle, a linear Kalman filter is used to complete the online estimation and correction of all error terms of the attitude chain based on the attitude observation data of the aircraft until all error terms of the corrected attitude chain are less than the set threshold, and the correction of all error terms of the attitude chain is completed, thereby realizing continuous correction of the aircraft attitude, improving the correction accuracy of the aircraft attitude, and meeting the requirements of high-precision attitude determination. Description of the Drawings

[0016] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the following will briefly introduce the drawings required for use in the embodiments. Obviously, the drawings in the following description are only some embodiments of the present application. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0017] Figure 1 It is a schematic flowchart of a method for continuously correcting the attitude of an aircraft based on a starlight navigation device provided by an embodiment of the present application; Figure 2Schematic structural diagram of an aircraft attitude continuous correction system based on a starlight navigation device provided by an embodiment of the present application.

[0018] Figure 3 Schematic diagram of the light source included angle measurement process provided by an embodiment of the present application; Figure 4 Schematic diagram of the star observation process provided by an embodiment of the present application.

[0019] Reference numerals: 11 - Outer frame, 12 - Inner frame, 13 - Starlight navigation device, 14 - Communication interface, 15 - Biaxial rotation mechanism, 16 - Navigation computer, 17 - Star sensor, 18 - Inner platform. Detailed implementation manners

[0020] Next, the technical solutions in the embodiments of the present application will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present application.

[0021] To make the above objects, features, and advantages of the present application more obvious and understandable, the present application will be further described in detail below in conjunction with the accompanying drawings and specific implementation manners.

[0022] In an exemplary embodiment, the present application provides an aircraft attitude continuous correction method based on a starlight navigation device. This method is executed by a computer device, which can be specifically executed by a computer device such as a terminal or a server alone, or jointly executed by a terminal and a server. In the embodiments of the present application, this method is described by taking it as an example applied to a server. As Figure 1 shown, this method includes: Step 100: In each iteration period, obtain the attitude observation data of the aircraft. The attitude observation data of the aircraft includes: inertial navigation installation reference vector, starlight navigation device installation reference vector, and the attitude matrix of the inertial navigation body system relative to the navigation system. One star observation process using the starlight navigation device is one iteration period. For example, the starlight navigation device can be used to perform the star observation process based on a static or quasi-static star observation strategy.

[0023] Step 101: Use a linear Kalman filter to complete the online estimation and correction of all error terms of the attitude chain based on the attitude observation data of the aircraft until the corrected all error terms of the attitude chain are less than a set threshold (for example, 5 arcseconds), and then complete the correction of all error terms of the attitude chain. The all error terms of the attitude chain include 15 error terms such as gyro error, platform error angle, and star-inertial installation error.

[0024] In another exemplary embodiment of the present application, based on the aircraft attitude continuous correction method provided above in the present application, it is assumed that 15 errors such as gyro error, platform error angle, and star-inertial installation error can be observed through 7 star observation processes (i.e., 7 iteration cycles), so as to establish a full-attitude link error state model. A linear Kalman filter is used as the optimal estimation method to complete the online estimation of all error terms of the attitude link and the attitude angle correction. Based on this, taking the platform error angle and the star-inertial installation error as an example of the specific iterative correction process, the implementation processes of steps 100 and 101 provided in the present application are described.

[0025] 1) The first star observation process (i.e., the first observation).

[0026] The star vector observation equation considering the star-inertial installation error is: (1) In the formula, is the inertial navigation installation reference vector of the first iteration cycle, is the starlight navigation device installation reference vector of 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 body system relative to the navigation system, is the star-inertial installation error defined in the inertial navigation body system of the first iteration cycle.

[0027] When calculating the platform error angle and the star-inertial installation error through a linear Kalman filter, the measurement equation of the linear Kalman filter is expressed as: (2) In the formula, is the right multiplication symbol.

[0028] The observation result obtained through the first star observation process can obtain the platform error angle and the star-inertial installation error of the first iteration cycle, and the attitude angle obtained by real-time calculation is corrected based on this, as follows: (3) In the formula, and are the correction value and the original value of the pitch angle of the first iteration cycle respectively, and are the correction value and the original value of the roll angle of the first iteration cycle respectively, and They are respectively the correction value and the original value of the heading angle in the first iteration cycle. 、 and They are respectively the components of the platform error angle vector in the navigation system along three axes. 、 and They are respectively the components of the star-inertial installation error in the navigation system along three axes, so as to compensate for the components of the platform error angle vector along three axes in the navigation system and the components of the star-inertial installation error along three axes in the navigation system Then judge whether the platform error angle and the star-inertial installation error are both less than 5 arcseconds. If so, complete the iteration; otherwise, proceed to the next iteration cycle.

[0029] 2) The second star observation process (i.e., the second observation).

[0030] The star vector observation equation considering the star-inertial installation error is: (4) In the formula, is the inertial navigation installation reference vector in the second iteration cycle, is the starlight navigation device installation reference vector in the second iteration cycle, is the platform error angle in the second iteration cycle, is the star-inertial installation error defined in the body frame of the inertial navigation in the second iteration cycle.

[0031] Next, calculate the platform error angle and the star-inertial installation error through a linear Kalman filter, and we have: (5) Through the observation results of the second star observation process, the platform error angle and the star-inertial installation error in the second iteration cycle can be obtained, and the attitude angle obtained by real-time calculation is corrected based on this, and we have: (6) In the formula, and They are respectively the correction value and the original value of the pitch angle in the second iteration cycle, and They are respectively the correction value and the original value of the roll angle in the second iteration cycle, and They are respectively the correction value and the original value of the heading angle in the second iteration cycle, 、 and They are respectively the components of the platform error angle vector in the navigation system along three axes in the second iteration cycle, 、 and They are respectively the components of the star-inertial installation error in the navigation system along three axes in the second iteration cycle.

[0032] Judge whether both the platform error angle and the star-inertial installation error are less than 5 arcseconds. If so, complete the iteration; otherwise, proceed to the next iteration cycle.

[0033] 3) The third star observation process (i.e., the third observation).

[0034] The star vector observation equation considering the star-inertial installation error is: (7) In the formula, is the inertial navigation installation reference vector in the third iteration cycle, is the starlight navigation device installation reference vector in the third iteration cycle, is the platform error angle in the third iteration cycle, is the star-inertial installation error defined in the inertial navigation body frame in the third iteration cycle.

[0035] Next, use the linear Kalman filter to calculate the platform error angle and the star-inertial installation error , and we have: (8) Through the results of the third observation, the platform error angle and the star-inertial installation error in the third iteration cycle can be obtained, and the attitude angle calculated in real time is corrected accordingly. We have: (9) In the formula, and are respectively the correction value and the original value of the pitch angle in the third iteration cycle, and are respectively the correction value and the original value of the roll angle in the third iteration cycle, and are respectively the correction value and the original value of the heading angle in the third iteration cycle, 、 and They are respectively the platform error angle vectors in the third iteration cycle in the navigation system and the components in three axial directions 、 and They are respectively the star-inertial installation errors in the third iteration cycle in the navigation system and the components in three axial directions

[0036] Then judge whether the platform error angle and the star-inertial installation error are both less than 5 arcseconds. If so, the iteration is completed; otherwise, proceed to the next iteration cycle

[0037] And so on, until the (k + 1)-th star observation process (i.e., the (k + 1)-th observation), there is The star vector observation equation considering the star-inertial installation error is shown in the following formula (10).

[0038] (10) In the formula is the inertial navigation installation reference vector in the (k + 1)-th iteration cycle is the starlight navigation device installation reference vector in the (k + 1)-th iteration cycle is the platform error angle in the (k + 1)-th iteration cycle is the star-inertial installation error defined in the body frame of the inertial navigation in the (k + 1)-th iteration cycle

[0039] Next, calculate the platform error angle and the star-inertial installation error through the linear Kalman filter, and there is (11) The platform error angle and the star-inertial installation error in the (k + 1)-th iteration cycle can be obtained from the observation results of the (k + 1)-th star observation process, and the attitude angles of real-time solution are corrected accordingly, and there is (12) In the formula and are respectively the correction value and the original value of the pitch angle in the (k + 1)-th iteration cycle and are respectively the correction value and the original value of the roll angle in the (k + 1)-th iteration cycle and are respectively the correction value and the original value of the heading angle in the (k + 1)-th iteration cycle 、 and They are respectively the platform error angle vectors in the (k + 1)-th iteration cycle in the navigation system and the components in three axial directions 、 and They are respectively the star-inertial installation errors in the (k + 1)-th iteration cycle in the navigation system and the components in three axial directions

[0040] Then, it is judged whether the platform error angle and the star-inertial installation error are both less than 5 arcseconds. If so, the iteration is completed; otherwise, the next iteration cycle is carried out

[0041] Based on the above description, in step 101, the process of correcting the attitude of the aircraft by using the corrected full attitude chain error term can be described as follows 1. Determine the components in three axial directions of the corrected star-inertial installation error in the navigation system

[0042] 2. Based on the components in three axial directions of the corrected star-inertial installation error in the navigation system and the corrected platform error angle, correct the attitude angles of the aircraft to complete the attitude correction of the aircraft. The attitude angles include: pitch angle, roll angle and heading angle

[0043] In summary, the present application corrects the full attitude chain error term of the aircraft system and continuously corrects the attitude of the aircraft, further improving the attitude accuracy of the aircraft, and can solve problems such as the deterioration of the combined accuracy caused by the inability of the aircraft system to model and decouple and identify the full system error term and the low attitude accuracy of the aircraft system

[0044] Based on the same inventive concept, the embodiment of the present application also provides an aircraft attitude continuous correction system based on a starlight navigation device for implementing the above-mentioned aircraft attitude continuous correction method. The implementation solutions provided by this system to solve problems are similar to the implementation solutions described in the above method. Therefore, the specific limitations in one or more embodiments of the aircraft attitude continuous correction system based on a starlight navigation device provided below can refer to the limitations on the aircraft attitude continuous correction method in the above text, and will not be repeated here

[0045] In an exemplary embodiment, as Figure 2 shown, an aircraft attitude continuous correction system based on a starlight navigation device is provided, including: a starlight navigation device 13, a two-axis indexing mechanism 15, a navigation computer 16, an optical theodolite (not shown in the figure), a first fixed light source (i.e., Figure 2 the light source 1 in Figure 2in the light source 2). Among them, the starlight navigation device 13 includes a star sensor 17.

[0046] The starlight navigation device 13 exchanges data with the navigation computer 16. The starlight navigation device 13 is installed on a two-axis indexing mechanism. The two-axis indexing mechanism 15 is installed on the airframe of the aircraft. The first fixed light source and the second fixed light source are arranged on the ground at a set angle.

[0047] The optical theodolite is used to provide a correction reference. For example, the true north direction is obtained by measuring the azimuth of the North Star with the optical theodolite and led to the prism on the pier, that is, the northward reference prism mirror surface for collimation. The northward reference prism is fixed on the pier after collimation to serve as the northward reference (i.e., the correction reference). Both the first fixed light source and the second fixed light source are used to simulate the starlight of the stars. The starlight navigation device 13, driven by the two-axis indexing mechanism 15, combines the first fixed light source, the second fixed light source and the correction reference to perform the star observation process. The navigation computer 16 is used to implement the aircraft attitude continuous correction method provided above based on the starlight navigation device to achieve continuous correction of the aircraft attitude.

[0048] As an alternative implementation, as Figure 2 shown, the starlight navigation device 13 exchanges data with the navigation computer 16 through the communication interface 14. Among them, the communication interface 14 can adopt a 1553b communication interface.

[0049] As an alternative implementation, during the actual installation process, as Figure 2 shown, the aircraft attitude continuous correction system provided by the present application can be installed inside the nose of the aircraft. The two-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.

[0050] In an exemplary embodiment, when implementing the continuous correction of the aircraft attitude using the aircraft attitude continuous correction system provided above in the present application, it is necessary to first decouple all error terms of the attitude chain and then perform continuous attitude correction. Based on this, in this embodiment, the implementation process of the continuous correction of the aircraft attitude includes: Step 1, install the system: Install the starlight navigation device 13 on the two-axis rotation mechanism 15, and the attitude adjustment of the starlight navigation device 13 can be realized through the two-axis rotation mechanism 15. Install two fixed light sources (i.e., the first fixed light source and the second fixed light source) on the stable base on the ground to simulate the real starlight of the stars and generate the star vectors 1 and 2 as Figure 2 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°. To ensure that the parallel 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 central height of the fixed light source is basically the same as the central height of the field of view of the star sensor 17 installed on the two-axis rotating mechanism 15, and the simulated stellar vector needs to be calibrated and measured before the test. The navigation computer 16 is connected to the starlight 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 by an optical theodolite to obtain the true north direction, and it is led to the mirror surface of the prism on the pier, that is, the northward reference prism, and collimated. After being collimated, the northward reference prism is fixed on the pier as the northward reference. The star sensor 17 corrects the attitude error and device error of the starlight navigation device 13 by measuring the known simulated starlight information.

[0051] Step 2: Establishment of the attitude reference of the fixed light source: Use an optical theodolite and the northward reference to directly measure the azimuth angle of the first fixed light source (i.e., the light source 1 in Figure 2 ). By adjusting the attitude angle of the indexing mechanism, the parallel beam of the first fixed light source just falls near the center point in the imaging plane of the star sensor 17 to ensure high-precision measurement of the starlight vector. Then rotate the outer frame 11 of the two-axis rotating mechanism 15 so that the parallel beam of the second fixed light source (i.e., the light source 2 in Figure 2 ) just falls at the previous imaging point in the imaging plane of the star sensor 17, and the rotation angle needs to be calculated by reading the output of the star sensor 17 in real time. When the star sensor 17 points to the first fixed light source, record the azimuth angle of the indexing mechanism as , and when it points to the second fixed light source, record the azimuth angle of the indexing mechanism as , so the included angle between the two fixed light sources can be expressed as , . And the azimuth angle of the second fixed light source is measured by measuring the azimuth angle of the first fixed light source and the included angle between the light sources. The process of measuring the included angle between the light sources is as shown in Figure 3 . Figure 3 In

[0052] Step 3: Analysis of the number of star observations: For the starlight navigation device 13, the total error terms of the attitude chain mainly include 15 error terms such as gyro zero position, gyro scale, gyro installation error, star-inertial installation error, and platform error angle. For the star sensor 17 with a small field of view, the observability of the star-inertial installation error in the direction of the optical axis is poor, so at least 7 star observations are required to make the total error terms observable.

[0053] Step 4, Stargazing process: Adopt a static or quasi-static stargazing strategy, that is, complete stargazing under the condition that the starlight navigation device 13 is completely static after attitude adjustment. The starlight navigation device 13 needs to perform initial alignment for 5 minutes first, and then carry out 7 stargazing operations successively as the indexing mechanism rotates. The specific stargazing process is as Figure 4 shown. Figure 4 In this case, the time measurement of light source 1 and light source 2, as well as the rotation angle values of the inner frame and the outer frame, are not specifically limited in this application and are determined according to measurement requirements during actual application.

[0054] Step 5, Correction of all error terms in the attitude chain: Through the above 7 stargazing processes, the observation of 15 errors such as gyro error, platform error angle, and star-inertial installation error can be completed, thus establishing a full attitude link error state model. Use a linear Kalman filter as the optimal estimation method to complete the online estimation and correction of all error terms in the attitude chain.

[0055] Step 6, Continuous attitude correction: Each observation is an iteration cycle. Repeat the measurement equation of the linear Kalman filter shown in formula (2), formula (5), formula (8), or formula (11) to obtain the platform error angle and the star-inertial installation error and compensate them to correct the attitude angle obtained by real-time calculation. The iteration process stops until both the platform error angle and the star-inertial installation error are less than 5 arcseconds, thereby realizing continuous attitude correction of the aircraft.

[0056] In an exemplary embodiment, a computer device is provided. The computer device can be a server or a terminal. The computer device includes a processor, a memory, an input / output interface (Input / Output, abbreviated as I / O), and a communication interface. Among them, the processor, the memory, and the input / output interface are connected through a system bus, and the communication interface is connected to the system bus through the input / output interface. Among them, the processor of the computer device is used 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 the computer program in the non-volatile storage medium. The database of the computer device is used to store aircraft attitude continuous correction data based on the starlight navigation device. The input / output interface of the computer device is used to exchange information between the processor and external devices. The communication interface of the computer device is used to communicate with external terminals through a network connection. When the computer program is executed by the processor, it realizes a method for continuous attitude correction of an aircraft based on a starlight navigation device.

[0057] In an exemplary embodiment, a computer device is provided, including a memory and a processor. A computer program is stored in the memory, and when the processor executes the computer program, the steps in the above method embodiments are implemented.

[0058] In an exemplary embodiment, a computer-readable storage medium is provided, storing a computer program, and when the computer program is executed by a processor, the steps in the above method embodiments are implemented.

[0059] In an exemplary embodiment, a computer program product is provided, including a computer program, and when the computer program is executed by a processor, the steps in the above method embodiments are implemented.

[0060] 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 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 need to comply with relevant regulations.

[0061] Those of ordinary skill in the art can understand that all or part of the processes of implementing the above method embodiments can be completed by instructing relevant hardware through a computer program. The computer program can be stored in a non-volatile computer-readable storage medium, and when the computer program is executed, it can include the processes of the above method embodiments. Among them, any reference to a memory, database, or other medium used in the embodiments provided in this application can include at least one of non-volatile and volatile memories. 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), magnetoresistive 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 can be in various forms, such as static random access memory (SRAM) or dynamic random access memory (DRAM), etc.

[0062] In each of the embodiments provided in this application, the database involved may include at least one of a relational database and a non-relational database. The non-relational database may include a distributed database based on blockchain, etc., without limitation. In each of the embodiments provided in this application, the processor may be a general-purpose processor, a central processing unit, a graphics processing unit, a digital signal processor, a programmable logic device, a data processing logic device based on quantum computing, etc., without limitation.

[0063] The technical features of the above embodiments can be combined arbitrarily. For the sake of brevity of description, 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, it should be considered as the scope recorded in this specification.

[0064] Specific examples are used in this article to elaborate on the principle and implementation manner of this application. The description of the above embodiments is only used to help understand the method and its core idea of this application; at the same time, for those of ordinary skill in the art, according to the idea of this application, there will be changes in the specific implementation manner and application scope. In summary, the content of this specification should not be construed as a limitation to this application.

Claims

1. A method for continuously correcting the attitude of an aircraft based on a starlight navigation device, characterized in that, including: In each iteration period, obtaining the attitude observation data of the aircraft; The attitude observation data of the aircraft includes: the inertial navigation installation reference vector, the starlight navigation device installation reference vector, and the attitude matrix of the inertial navigation body system relative to the navigation system; One star observation process using the starlight navigation device is one iteration period; Using a linear Kalman filter to complete the online estimation and correction of the full attitude chain error term based on the attitude observation data of the aircraft. When the corrected full attitude chain error term is less than the set threshold, the correction of the full attitude chain error term is completed.

2. The method for continuously correcting the attitude of an aircraft based on a starlight navigation device according to claim 1, wherein The full attitude chain error term includes: gyro error, platform error angle, and star-inertial installation error.

3. The method for continuously correcting the attitude of an aircraft based on a starlight navigation device according to claim 1, characterized in that, Using the starlight navigation device to perform the star observation process based on a static or quasi-static star observation strategy.

4. The method for continuously correcting the attitude of an aircraft based on a starlight navigation device according to claim 1, wherein The measurement equation of the linear Kalman filter is expressed as: ； In the formula, is the platform error angle in the (k + 1)-th iteration cycle, is the star-inertial installation error in the (k + 1)-th iteration cycle, is the inertial navigation installation reference vector in the (k + 1)-th iteration cycle, is the starlight navigation device installation reference vector in the (k + 1)-th iteration cycle, is the identity matrix, is the attitude matrix of the inertial navigation body frame relative to the navigation frame, is the right multiplication symbol.

5. The method for continuously correcting the attitude of an aircraft based on a starlight navigation device according to claim 1, wherein The maximum value of the set threshold is 5 arcseconds.

6. The method for continuously correcting the attitude of an aircraft based on a starlight navigation device according to claim 2, characterized in that Using the corrected full attitude chain error term to correct the attitude of the aircraft, including: Determining the components of the corrected star-inertial installation error in the three axes of the navigation system; Based on the components of the corrected star-inertial installation error in the three axes of the navigation system and the corrected platform error angle, correcting the attitude angle of the aircraft to complete the attitude correction of the aircraft; The attitude angles include: pitch angle, roll angle, and heading angle.

7. The method for continuously correcting the attitude of an aircraft based on a starlight navigation device according to claim 1, characterized in that, The number of iteration periods is at least 7.

8. An aircraft attitude continuous correction system based on a starlight navigation device, characterized in that, The aircraft attitude continuous correction system based on the starlight navigation device includes: a starlight navigation device, a two-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 installed on the two-axis indexing mechanism; The two-axis indexing mechanism is installed on the airframe 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; Both the first fixed light source and the second fixed light source are used to simulate the starlight of a star; The starlight navigation device, driven by the two-axis indexing mechanism, combines the first fixed light source, the second fixed light source, and the correction reference to perform the star observation process; The navigation computer is used to implement the aircraft attitude continuous correction method based on the starlight navigation device as described in any one of claims 1-7 to achieve continuous correction of the aircraft attitude.

9. The aircraft attitude continuous correction system based on the starlight navigation device according to claim 8, wherein, The starlight navigation device includes a star sensor.

10. The aircraft attitude continuous correction system based on a starlight navigation device according to claim 8, characterized in that, The aircraft attitude continuous correction system based on the starlight navigation device further includes: a communication interface; The starlight navigation device exchanges data with the navigation computer through the communication interface.

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