A small field of view celestial navigation method and device based on asynchronous multi-star vector positioning

Abstract

This invention provides a small field-of-view astronomical navigation method and device based on asynchronous multi-satellite vector attitude determination. The method includes: projecting multi-satellite observation vectors acquired at different observation times onto the device coordinate system at the last observation time; calculating a first attitude matrix of the device coordinate system relative to the inertial coordinate system based on the projected multi-satellite observation vectors and a reference vector; calculating the geographical location of the vehicle based on the first attitude matrix and a second attitude matrix of the vehicle relative to the geographic coordinate system; and correcting the inertial navigation accumulated error based on the first attitude matrix and the geographical location. Compared with the traditional altitude difference method, this invention effectively suppresses the impact of inertial navigation error on navigation accuracy and no longer restricts the vehicle's motion state during the observation period, maintaining high navigation accuracy even in non-stationary or non-uniform straight-line states. Since the calculated parameters can obtain the vehicle's geographical location and three-axis attitude information relative to the inertial coordinate system, it can provide more comprehensive reference navigation parameters.

Description

Technical Field

[0001] This invention belongs to the field of navigation, and more specifically, relates to a small field-of-view astronomical navigation method and device based on asynchronous multi-star vector attitude determination. Background Technology

[0002] Near-surface carriers operate within the atmosphere, where background noise significantly impacts astronomical observations. Therefore, astronomical navigation primarily employs high-resolution, small-field-of-view astronomical navigation systems.

[0003] The "altitude difference method," a classic small-field-of-view astronomical navigation algorithm, is the primary algorithm used in astronomical navigation systems for ships, submarines, and other vessels. Its navigation process relies on inertial navigation (INS). First, the INS obtains the initial position and attitude of the vessel to track and search for the observed celestial body. Second, it uses the horizontal reference provided by the INS to obtain the observed altitude of the celestial body. Then, it calculates the range and star angle based on the INS information to correct for eccentricity. Finally, it uses its own calculated position and heading to correct the accumulated errors of the INS.

[0004] It can be seen that the solution process of the "altitude difference method" inevitably introduces inertial navigation error, which affects navigation accuracy. At the same time, in order to ensure the accuracy of the eccentricity difference correction, the vehicle must remain stationary or in a uniform straight-line state during the observation period. The navigation accuracy is poor when the vehicle is not stationary or in a non-uniform straight-line state. Summary of the Invention

[0005] To address the shortcomings of existing technologies, the present invention aims to provide a small field-of-view astronomical navigation method and device based on asynchronous multi-star vector attitude determination, which aims to solve the problem of inertial navigation error affecting navigation accuracy during the "altitude difference method" calculation process, as well as the problem of poor navigation accuracy in non-stationary or non-uniform speed straight flight states.

[0006] In a first aspect, the present invention provides a small field-of-view astronomical navigation method based on asynchronous multi-star vector attitude determination, comprising:

[0007] Project the multi-star observation vectors acquired at different observation times onto the device coordinate system at the last observation time;

[0008] Based on the projected multi-satellite observation vector and reference vector, the first attitude matrix of the equipment coordinate system relative to the inertial coordinate system is calculated;

[0009] Based on the first attitude matrix and the second attitude matrix of the vehicle relative to the geographic coordinate system output by the inertial navigation system, the geographical location of the vehicle is calculated;

[0010] Based on the first attitude matrix and the geographical location, the cumulative error of the inertial navigation system is corrected.

[0011] In some embodiments, projecting the multi-star observation vectors acquired at different observation times onto the device coordinate system at the last observation time includes:

[0012] Based on the time matrix, position matrix, and second attitude matrix of the target observation time, the observation vector of the target observed star at the target observation time is projected into the device coordinate system at the last observation time.

[0013] In some embodiments, the projection of the observation vector of the target observed star at the target observation time onto the device coordinate system at the last observation time satisfies the following calculation formula:

[0014]

[0015] Where i, e, n, b, and s represent the inertial coordinate system, Earth coordinate system, geographic coordinate system, carrier coordinate system, and equipment coordinate system, respectively; m represents the number of observed stars; j represents the index of the target observed star; and T... j T represents the target observation time. m Indicates the last observation time. Indicates the target observed star at T j The observation vector in the device coordinate system s at the observation time. Indicates the target observed star at T m The observation vector in the device coordinate system s at the observation time. and T represents respectively j The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, and T represents respectively m The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, This represents the installation error matrix from the carrier coordinate system b to the equipment coordinate system s. The superscript T indicates the transpose operator, and the superscript ~ indicates the observation containing the error in the corresponding variable.

[0016] In some embodiments, calculating the first attitude matrix of the device coordinate system relative to the inertial coordinate system based on the projected multi-satellite observation vector and the reference vector includes:

[0017] Based on the projected multi-star observation vector and the reference vector, an optimal estimation model is established;

[0018] Solve the optimal estimation model to determine the first attitude matrix of the device coordinate system relative to the inertial coordinate system.

[0019] In some embodiments, the method further includes:

[0020] Based on the first attitude matrix and the installation error matrix from the carrier coordinate system to the device coordinate system, the third attitude matrix of the carrier relative to the inertial coordinate system is determined.

[0021] In some embodiments, calculating the geographical location of the vehicle based on the first attitude matrix and the second attitude matrix of the vehicle relative to the geographic coordinate system output by the inertial navigation system includes:

[0022] Based on the first attitude matrix, the installation error matrix from the carrier coordinate system to the device coordinate system, the time matrix at the last observation time, and the second attitude matrix, the geographical location of the carrier is calculated.

[0023] In some embodiments, correcting the cumulative error of the inertial navigation system based on the first attitude matrix and the geographical location includes:

[0024] Based on the first attitude matrix and the geographical location, attitude measurement equations and position measurement equations are established;

[0025] Based on the attitude measurement equation and the position measurement equation, the cumulative error of the inertial navigation system is corrected by integrated navigation.

[0026] Secondly, the present invention provides a small field-of-view astronomical navigation device based on asynchronous multi-star vector attitude determination, comprising:

[0027] The projection unit is used to project the multi-star observation vectors acquired at different observation times onto the device coordinate system at the last observation time;

[0028] The first calculation unit is used to calculate the first attitude matrix of the device coordinate system relative to the inertial coordinate system based on the projected multi-star observation vector and the reference vector.

[0029] The second calculation unit is used to calculate the geographical location of the vehicle based on the first attitude matrix and the second attitude matrix of the vehicle relative to the geographic coordinate system output by the inertial navigation system.

[0030] The correction unit is used to correct the cumulative error of the inertial navigation system based on the first attitude matrix and the geographical location.

[0031] Thirdly, the present invention provides an electronic device comprising: at least one memory for storing a program; and at least one processor for executing the program stored in the memory, wherein when the program stored in the memory is executed, the processor is configured to execute the method described in the first aspect or any possible implementation thereof.

[0032] Fourthly, the present invention provides a computer-readable storage medium storing a computer program that, when run on a processor, causes the processor to perform the method described in the first aspect or any possible implementation thereof.

[0033] Fifthly, the present invention provides a computer program product that, when run on a processor, causes the processor to perform the method described in the first aspect or any possible implementation thereof.

[0034] The present invention provides a small field-of-view astronomical navigation method and device based on asynchronous multi-satellite vector attitude determination. This method projects multi-satellite observation vectors acquired at different observation times onto the device coordinate system at the last observation time, correcting and unifying the multi-satellite observation vectors acquired at different observation times. Then, using the projected multi-satellite observation vectors and a reference vector, the first attitude matrix of the device coordinate system relative to the inertial coordinate system is calculated. Using the first attitude matrix and the second attitude matrix of the carrier output by the inertial navigation system relative to the geographic coordinate system, the geographical location of the carrier is calculated. Finally, the cumulative error of the inertial navigation system is corrected using the carrier's geographical location and the first attitude matrix relative to the inertial coordinate system. Compared with the traditional altitude difference method, this method effectively suppresses the impact of inertial navigation errors on navigation accuracy and no longer restricts the carrier's motion state during the observation period. Even in non-stationary or non-uniform straight-line states, high navigation accuracy can be maintained. Furthermore, since the calculated parameters can obtain the carrier's geographical location and three-axis attitude information relative to the inertial coordinate system, more comprehensive reference navigation parameters can be provided. Attached Figure Description

[0035] To more clearly illustrate the technical solutions in this invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of this invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0036] Figure 1 This is a flowchart illustrating the small field-of-view astronomical navigation method based on asynchronous multi-star vector attitude determination provided in an embodiment of the present invention;

[0037] Figure 2 This is a schematic diagram of the structure of a small field-of-view astronomical navigation device based on asynchronous multi-star vector attitude determination provided in an embodiment of the present invention;

[0038] Figure 3 This is a schematic diagram of the structure of the electronic device provided in an embodiment of the present invention. Detailed Implementation

[0039] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the invention.

[0040] Figure 1 This is a flowchart illustrating the small field-of-view astronomical navigation method based on asynchronous multi-star vector attitude determination provided in an embodiment of the present invention, as shown below. Figure 1 As shown, the method includes at least the following steps:

[0041] S101. Project the multi-star observation vectors acquired at different observation times onto the device coordinate system at the last observation time.

[0042] Specifically, the observation information directly obtained by the small field-of-view astronomical navigation system is the observation vector of the observed star in the device coordinate system. At the same time, the observation method of tracking the star one by one is adopted, which means that the observation vector of each observed star is not projected into the device coordinate system at the same observation time, and must be corrected and unified.

[0043] It is conceivable that by projecting the multi-satellite observation vectors acquired at different observation times onto the device coordinate system at any specified observation time, the unification of the multi-satellite observation vectors acquired at different observation times can be achieved. However, due to the need for real-time acquisition of navigation information, and considering the feasibility of operation, it is preferable to project the multi-satellite observation vectors acquired at different observation times onto the device coordinate system at the last observation time.

[0044] S102. Based on the projected multi-satellite observation vector and reference vector, calculate the first attitude matrix of the equipment coordinate system relative to the inertial coordinate system.

[0045] Specifically, a certain transformation relationship exists between the observed vector of the observed star and the reference vector, and the reference vector is calculated from the apparent position of the celestial body. After obtaining the multi-star observed vector projected in S101, the first attitude matrix of the device coordinate system relative to the inertial coordinate system is calculated by combining it with the reference vector of the observed star.

[0046] Optionally, after calculating the first attitude matrix of the device coordinate system relative to the inertial coordinate system, the attitude matrix of the carrier relative to the inertial coordinate system can be further calculated by combining the installation error matrix from the carrier coordinate system to the device coordinate system, and used as the reference navigation parameter.

[0047] S103. Based on the first attitude matrix of the device coordinate system relative to the inertial coordinate system and the second attitude matrix of the carrier relative to the geographic coordinate system output by the inertial navigation system, calculate the geographical location of the carrier.

[0048] Specifically, after obtaining the first attitude matrix of the device coordinate system relative to the inertial coordinate system in S102, the second attitude matrix of the carrier relative to the geographic coordinate system output by the inertial navigation system is combined with the relevant formulas to calculate the geographical location of the carrier.

[0049] S104. Based on the first attitude matrix of the device coordinate system relative to the inertial coordinate system and the geographical location, the cumulative error of the inertial navigation system is corrected.

[0050] Specifically, after obtaining the first attitude matrix of the device coordinate system relative to the inertial coordinate system and the geographical location of the carrier, measurement equations can be established, and the cumulative error of the inertial navigation system can be corrected through integrated navigation.

[0051] The small field-of-view astronomical navigation method based on asynchronous multi-satellite vector attitude determination provided in this invention projects multi-satellite observation vectors acquired at different observation times onto the device coordinate system at the last observation time, thereby correcting and unifying the multi-satellite observation vectors acquired at different observation times. Then, using the projected multi-satellite observation vectors and the reference vector, the first attitude matrix of the device coordinate system relative to the inertial coordinate system is calculated. Using the first attitude matrix and the second attitude matrix of the vehicle output by the inertial navigation system relative to the geographic coordinate system, the geographical location of the vehicle is calculated. Finally, the cumulative error of the inertial navigation system is corrected using the geographical location of the vehicle and the first attitude matrix relative to the inertial coordinate system. Compared with the traditional altitude difference method, this method effectively suppresses the impact of inertial navigation error on navigation accuracy. At the same time, due to the correction and unification of the observation vectors, the motion state of the vehicle during the observation period is no longer restricted. Even in non-stationary or non-uniform straight-line states, high navigation accuracy can be maintained. Furthermore, the geographical location of the vehicle and its three-axis attitude information relative to the inertial coordinate system are calculated, providing more comprehensive reference navigation parameters.

[0052] In some embodiments, S101 specifically includes:

[0053] Based on the time matrix, position matrix, and second attitude matrix of the target observation time, and the time matrix, position matrix, and second attitude matrix of the last observation time, the observation vector of the target star at the target observation time is projected into the device coordinate system at the last observation time; wherein, the position matrix and the second attitude matrix are output by the inertial navigation system.

[0054] Specifically, in the process of correcting asynchronous multi-satellite vectors, projecting the multi-satellite observation vectors obtained at different observation times onto the device coordinate system at the last observation time is achieved based on the position matrix and second attitude matrix output by the inertial navigation system, as well as the time matrix of the observation time.

[0055] Optionally, the observation vector of the target star at the target observation time is projected onto the device coordinate system at the last observation time, satisfying the following calculation formula:

[0056]

[0057] Where i, e, n, b, and s represent the inertial coordinate system, Earth coordinate system, geographic coordinate system, carrier coordinate system, and equipment coordinate system, respectively; m represents the number of observed stars; j represents the index of the target observed star; and T... j T represents the target observation time. m Indicates the last observation time. Indicates the target observed star at T j The observation vector in the device coordinate system s at the observation time. Indicates the target observed star at T m The observation vector in the device coordinate system s at the observation time. and T represents respectively j The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, and T represents respectively m The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, This represents the installation error matrix from the carrier coordinate system b to the equipment coordinate system s. The superscript T indicates the transpose operator, and the superscript ~ indicates the observation containing the error in the corresponding variable.

[0058] It can be seen that in the process of asynchronous multi-star vector correction, the time matrix of the observation time, as well as the position matrix and second attitude matrix output by the inertial navigation system are used in the above projection process. However, for the overall coordinate transformation process, it can be regarded as the effect of attitude rotation modulation and position difference. Therefore, it effectively suppresses the influence of inertial navigation error on coordinate transformation accuracy. Compared with the traditional "altitude difference method", the small field of view astronomical navigation method based on asynchronous multi-star vector attitude determination provided by this invention has obvious advantages in the correction of asynchronous observation data.

[0059] In some embodiments, S102 specifically includes:

[0060] An optimal estimation model is established based on the projected multi-satellite observation vector and the baseline vector.

[0061] Solve the optimal estimation model to determine the first attitude matrix of the equipment coordinate system relative to the inertial coordinate system.

[0062] Specifically, the observation vector and the reference vector of the observed star satisfy a certain transformation relationship, which is the multi-star vector attitude determination formula, satisfying:

[0063]

[0064] in, This represents the observation vector of the observed star after projection in S101; The reference vector representing the observed star is calculated from the apparent position of the celestial body; T represents m The first attitude matrix of the device coordinate system s relative to the inertial coordinate system i at the observation time.

[0065] When the number of observed stars is greater than or equal to 3, the attitude matrix can be solved using an optimal estimation algorithm. Based on the projected multi-satellite observation vector and the baseline vector, the optimal estimation model is established as follows:

[0066]

[0067] Where J() represents the optimal estimation function, w j This represents the weighting coefficient.

[0068] Thus, the three-axis attitude information of the carrier was obtained, which serves as the reference navigation parameters.

[0069] In some embodiments, after obtaining the first attitude matrix of the device coordinate system relative to the inertial coordinate system... Then, the third attitude matrix of the carrier relative to the inertial coordinate system was obtained. Specifically, this can be achieved through the installation error matrix between the carrier coordinate system and the equipment coordinate system. It is known that the observation equipment is mounted on the carrier, and the installation error angle between the two is fixed, i.e., the installation error matrix. Fixed and known in advance.

[0070] In some embodiments, S103 specifically includes:

[0071] Based on the first attitude matrix of the equipment coordinate system relative to the inertial coordinate system, the installation error matrix of the carrier coordinate system to the equipment coordinate system, the time matrix at the last observation time, and the second attitude matrix of the carrier relative to the geographic coordinate system output by the inertial navigation system, the geographical location of the carrier is calculated.

[0072] Specifically, the formula for calculating the carrier position satisfies:

[0073]

[0074] in, Indicates the geographical location of the carrier. The time matrix representing the last observation time, This represents the first attitude matrix of the device coordinate system relative to the inertial coordinate system. This represents the installation error matrix from the carrier coordinate system b to the equipment coordinate system s. This represents the second attitude matrix of the carrier relative to the geographic coordinate system, output by the inertial navigation system at the last observation moment.

[0075] Furthermore, the geographical location of the carrier satisfy:

[0076]

[0077] in, and These represent the latitude and longitude of the carrier, respectively.

[0078] Thus, the geographical location of the vehicle was determined and used as the reference navigation parameter.

[0079] In some embodiments, S104 specifically includes:

[0080] Based on the first attitude matrix of the equipment coordinate system relative to the inertial coordinate system and the geographical location of the carrier, attitude measurement equations and position measurement equations are established.

[0081] Based on attitude measurement equations and position measurement equations, the cumulative error of inertial navigation is corrected through integrated navigation.

[0082] Specifically, after obtaining the first pose matrix and the latitude of the carrier and longitude Then, the measurement equation can be established.

[0083] The attitude measurement equation is as follows:

[0084]

[0085] Where m2rv() represents the matrix to equivalent rotated vector transformation function, φ represents the misalignment angle, μ represents the equipment installation error angle, and δP=[δL δλ δh] T Let N represent the position error vector, with each component representing latitude error, longitude error, and elevation error, respectively.

[0086]

[0087] The position measurement equation is as follows:

[0088]

[0089] in, and This is the latitude and longitude output by the inertial navigation system at the last observation moment.

[0090] After establishing the attitude measurement equations and position measurement equations, the cumulative error of the inertial navigation system is corrected through integrated navigation.

[0091] Figure 2 This is a schematic diagram of the structure of a small field-of-view astronomical navigation device based on asynchronous multi-star vector attitude determination provided in an embodiment of the present invention, as shown below. Figure 2 As shown, the device includes at least:

[0092] Projection unit 201 is used to project the multi-star observation vectors acquired at different observation times onto the device coordinate system at the last observation time;

[0093] The first calculation unit 202 is used to calculate the first attitude matrix of the device coordinate system relative to the inertial coordinate system based on the projected multi-star observation vector and the reference vector.

[0094] The second calculation unit 203 is used to calculate the geographical location of the vehicle based on the first attitude matrix and the second attitude matrix of the vehicle relative to the geographic coordinate system output by the inertial navigation system.

[0095] The correction unit 204 is used to correct the cumulative error of the inertial navigation system based on the first attitude matrix and the geographical location.

[0096] In some embodiments, the projection unit 201 includes:

[0097] The projection module is used to project the observation vector of the target observed star at the target observation time onto the device coordinate system at the last observation time, based on the time matrix, position matrix, and second attitude matrix of the target observation time, and the time matrix, position matrix, and second attitude matrix of the last observation time; wherein the position matrix and the second attitude matrix are output by the inertial navigation system.

[0098] In some embodiments, the projection of the observation vector of the target observed star at the target observation time onto the device coordinate system at the last observation time satisfies the following calculation formula:

[0099]

[0100] Where i, e, n, b, and s represent the inertial coordinate system, Earth coordinate system, geographic coordinate system, carrier coordinate system, and equipment coordinate system, respectively; m represents the number of observed stars; j represents the index of the target observed star; and T... j T represents the target observation time. m Indicates the last observation time. Indicates the target observed star at T j The observation vector in the device coordinate system s at the observation time. Indicates the target observed star at T m The observation vector in the device coordinate system s at the observation time. and T represents respectivelyj The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, and T represents respectively m The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, This represents the installation error matrix from the carrier coordinate system b to the equipment coordinate system s. The superscript T indicates the transpose operator, and the superscript ~ indicates the observation containing the error in the corresponding variable.

[0101] In some embodiments, the first calculation unit 201 includes:

[0102] The first establishment module is used to establish an optimal estimation model based on the projected multi-star observation vector and the reference vector;

[0103] The solution module is used to solve the optimal estimation model and determine the first attitude matrix.

[0104] In some embodiments, the apparatus further includes:

[0105] The determining unit is used to determine the third attitude matrix of the carrier relative to the inertial coordinate system based on the first attitude matrix and the installation error matrix from the carrier coordinate system to the device coordinate system.

[0106] In some embodiments, the second calculation unit 202 includes:

[0107] The calculation module is used to calculate the geographical location of the carrier based on the first attitude matrix, the installation error matrix from the carrier coordinate system to the device coordinate system, the time matrix at the last observation time, and the second attitude matrix.

[0108] In some embodiments, the correction unit 204 includes:

[0109] The second establishment module is used to establish attitude measurement equations and position measurement equations based on the first attitude matrix and the geographical location;

[0110] The correction module is used to correct the cumulative error of the inertial navigation system through integrated navigation based on the attitude measurement equation and the position measurement equation.

[0111] It is understood that the detailed functional implementation of each of the above units / modules can be found in the description in the aforementioned method embodiments, and will not be repeated here.

[0112] It should be understood that the above-described device is used to execute the methods in the above embodiments. The implementation principle and technical effect of the corresponding program modules in the device are similar to those described in the above methods. The working process of the device can be referred to the corresponding process in the above methods, and will not be repeated here.

[0113] Based on the methods described in the above embodiments, this invention provides an electronic device. The device may include at least one memory for storing a program and at least one processor for executing the program stored in the memory. When the program stored in the memory is executed, the processor performs the methods described in the above embodiments.

[0114] Figure 3 This is a schematic diagram of the structure of the electronic device provided in the embodiment of the present invention, such as... Figure 3 As shown, the electronic device may include a processor 301, a communications interface 320, a memory 303, and a communication bus 304. The processor 301, communications interface 302, and memory 303 communicate with each other via the communication bus 304. The processor 301 can call software instructions stored in the memory 303 to execute the methods described in the above embodiments.

[0115] Based on the methods in the above embodiments, this embodiment of the invention provides a computer-readable storage medium storing a computer program that, when run on a processor, causes the processor to execute the methods in the above embodiments.

[0116] Based on the methods in the above embodiments, this embodiment of the invention provides a computer program product that, when run on a processor, causes the processor to execute the methods in the above embodiments.

[0117] It is understood that the processor in the embodiments of the present invention can be a Central Processing Unit (CPU), or other general-purpose processors, digital signal processors (DSPs), application-specific integrated circuits (ASICs), field-programmable gate arrays (FPGAs), or other programmable logic devices, transistor logic devices, hardware components, or any combination thereof. The general-purpose processor can be a microprocessor or any conventional processor.

[0118] The method steps in these embodiments of the invention can be implemented in hardware or by a processor executing software instructions. The software instructions can consist of corresponding software modules, which can be stored in random access memory (RAM), flash memory, read-only memory (ROM), programmable read-only memory (PROM), erasable programmable read-only memory (EPROM), electrically erasable programmable read-only memory (EEPROM), registers, hard disks, portable hard disks, CD-ROMs, or any other form of storage medium known in the art. An exemplary storage medium is coupled to the processor, enabling the processor to read information from and write information to the storage medium. Of course, the storage medium can also be a component of the processor. The processor and the storage medium can reside in an ASIC.

[0119] In the above embodiments, implementation can be achieved entirely or partially through software, hardware, firmware, or any combination thereof. When implemented using software, it can be implemented entirely or partially as a computer program product. The computer program product includes one or more computer instructions. When the computer program instructions are loaded and executed on a computer, all or part of the processes or functions described in the embodiments of the present invention are generated. The computer can be a general-purpose computer, a special-purpose computer, a computer network, or other programmable device. The computer instructions can be stored in a computer-readable storage medium or transmitted through the computer-readable storage medium. The computer instructions can be transmitted from one website, computer, server, or data center to another website, computer, server, or data center via wired (e.g., coaxial cable, fiber optic, digital subscriber line (DSL)) or wireless (e.g., infrared, wireless, microwave, etc.) means. The computer-readable storage medium can be any available medium that a computer can access or a data storage device such as a server or data center that integrates one or more available media. The available medium can be a magnetic medium (e.g., floppy disk, hard disk, magnetic tape), an optical medium (e.g., DVD), or a semiconductor medium (e.g., solid-state disk (SSD)).

[0120] It is understood that the various numerical designations used in the embodiments of the present invention are merely for the convenience of description and are not intended to limit the scope of the embodiments of the present invention.

[0121] Those skilled in the art will readily understand that the above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A small field-of-view astronomical navigation method based on asynchronous multi-star vector attitude determination, characterized in that, include: Project the multi-star observation vectors acquired at different observation times onto the device coordinate system at the last observation time; Based on the projected multi-satellite observation vector and reference vector, the first attitude matrix of the equipment coordinate system relative to the inertial coordinate system is calculated; Based on the first attitude matrix and the second attitude matrix of the vehicle relative to the geographic coordinate system output by the inertial navigation system, the geographical location of the vehicle is calculated; Based on the first attitude matrix and the geographical location, the cumulative error of the inertial navigation system is corrected; The step of projecting the multi-star observation vectors acquired at different observation times onto the device coordinate system at the last observation time includes: Based on the time matrix, position matrix, and second attitude matrix of the target observation time, the time matrix, position matrix, and second attitude matrix of the last observation time are used to project the observation vector of the target observed star at the target observation time into the device coordinate system at the last observation time. The projection of the observation vector of the target observed star at the target observation time onto the device coordinate system at the last observation time satisfies the following calculation formula: Where i, e, n, b, and s represent the inertial coordinate system, the Earth coordinate system, the geographic coordinate system, the carrier coordinate system, and the equipment coordinate system, respectively; m represents the number of observed stars; and j represents the index of the target observed star. Indicates the time of target observation. Indicates the last observation time. Indicates the target observed star at The observation vector in the device coordinate system s at the observation time. Indicates the target observed star at The observation vector in the device coordinate system s at the observation time. , and They represent The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, , and They represent The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, This represents the installation error matrix from the carrier coordinate system b to the equipment coordinate system s. The superscript T indicates the transpose operator, and the superscript ~ indicates that the corresponding variable contains an observation with error. The calculation of the first attitude matrix of the device coordinate system relative to the inertial coordinate system based on the projected multi-satellite observation vector and reference vector includes: Based on the projected multi-star observation vector and the reference vector, an optimal estimation model is established; Solve the optimal estimation model to determine the first attitude matrix; The observation vector and reference vector of the observed star satisfy the multi-star vector attitude determination formula: in, The observation vector representing the observed star after projection; The reference vector representing the observed star is calculated from the apparent position of the celestial body; express The first attitude matrix of the device coordinate system s relative to the inertial coordinate system i at the observation time; When the number of observed stars is greater than or equal to 3, the attitude matrix is ​​solved using the optimal estimation algorithm. Based on the projected multi-satellite observation vectors and the reference vector, the optimal estimation model is established as follows: in, Describes the optimal estimation function. This represents the weighting coefficient.

2. The small field-of-view astronomical navigation method based on asynchronous multi-star vector attitude determination according to claim 1, characterized in that, The method further includes: Based on the first attitude matrix and the installation error matrix from the carrier coordinate system to the device coordinate system, the third attitude matrix of the carrier relative to the inertial coordinate system is determined.

3. The small field-of-view astronomical navigation method based on asynchronous multi-star vector attitude determination according to claim 1, characterized in that, The calculation of the geographical location of the vehicle based on the first attitude matrix and the second attitude matrix of the vehicle relative to the geographic coordinate system output by the inertial navigation system includes: Based on the first attitude matrix, the installation error matrix from the carrier coordinate system to the device coordinate system, the time matrix at the last observation time, and the second attitude matrix, the geographical location of the carrier is calculated.

4. The small field-of-view astronomical navigation method based on asynchronous multi-star vector attitude determination according to claim 1, characterized in that, The step of correcting the cumulative error of the inertial navigation system based on the first attitude matrix and the geographical location includes: Based on the first attitude matrix and the geographical location, attitude measurement equations and position measurement equations are established; Based on the attitude measurement equation and the position measurement equation, the cumulative error of the inertial navigation system is corrected by integrated navigation.

5. A small field-of-view astronomical navigation device based on asynchronous multi-star vector attitude determination, characterized in that, include: The projection unit is used to project the multi-star observation vectors acquired at different observation times onto the device coordinate system at the last observation time; The first calculation unit is used to calculate the first attitude matrix of the device coordinate system relative to the inertial coordinate system based on the projected multi-star observation vector and the reference vector. The second calculation unit is used to calculate the geographical location of the vehicle based on the first attitude matrix and the second attitude matrix of the vehicle relative to the geographic coordinate system output by the inertial navigation system. A correction unit is used to correct the cumulative error of the inertial navigation system based on the first attitude matrix and the geographical location. The step of projecting the multi-star observation vectors acquired at different observation times onto the device coordinate system at the last observation time includes: Based on the time matrix, position matrix, and second attitude matrix of the target observation time, the time matrix, position matrix, and second attitude matrix of the last observation time are used to project the observation vector of the target observed star at the target observation time into the device coordinate system at the last observation time. The projection of the observation vector of the target observed star at the target observation time onto the device coordinate system at the last observation time satisfies the following calculation formula: Where i, e, n, b, and s represent the inertial coordinate system, the Earth coordinate system, the geographic coordinate system, the carrier coordinate system, and the equipment coordinate system, respectively; m represents the number of observed stars; and j represents the index of the target observed star. Indicates the time of target observation. Indicates the last observation time. Indicates the target observed star at The observation vector in the device coordinate system s at the observation time. Indicates the target observed star at The observation vector in the device coordinate system s at the observation time. , and They represent The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, , and They represent The time matrix at the observation time, the position matrix and the second attitude matrix output by the inertial navigation system, This represents the installation error matrix from the carrier coordinate system b to the equipment coordinate system s. The superscript T indicates the transpose operator, and the superscript ~ indicates that the corresponding variable contains an observation with error. The calculation of the first attitude matrix of the device coordinate system relative to the inertial coordinate system based on the projected multi-satellite observation vector and reference vector includes: Based on the projected multi-star observation vector and the reference vector, an optimal estimation model is established; Solve the optimal estimation model to determine the first attitude matrix; The observation vector and reference vector of the observed star satisfy the multi-star vector attitude determination formula: in, The observation vector representing the observed star after projection; The reference vector representing the observed star is calculated from the apparent position of the celestial body; express The first attitude matrix of the device coordinate system s relative to the inertial coordinate system i at the observation time; When the number of observed stars is greater than or equal to 3, the attitude matrix is ​​solved using the optimal estimation algorithm. Based on the projected multi-satellite observation vectors and the reference vector, the optimal estimation model is established as follows: in, Describes the optimal estimation function. This represents the weighting coefficient.

6. An electronic device, characterized in that, include: At least one memory for storing programs; At least one processor is configured to execute a program stored in the memory, wherein when the program stored in the memory is executed, the processor is configured to perform the method as described in any one of claims 1-4.

7. A computer-readable storage medium storing a computer program, characterized in that, When the computer program is run on the processor, it causes the processor to perform the method as described in any one of claims 1-4.

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