Le group-based inertial navigation / radar / satellite navigation three-combination transmitting system integrated navigation method

Abstract

The invention particularly relates to an inertial navigation / radar / satellite navigation three-combination emitter integrated navigation method based on Lie group, which comprises the following steps: responding to a navigation control instruction, and configuring a navigation system of an aircraft to enter an integrated navigation state; constructing a special European group based on the attitude features, the speed features and the position features, and establishing a motion description model based on the description of the special European group; acquiring navigation information of an inertial navigation system, and resolving the navigation information by using an error model under a special European group to obtain a state matrix; obtaining current measurement information; wherein the measurement information comprises satellite navigation subsystem measurement information and / or radar navigation subsystem measurement information; based on the measurement information, establishing an observation vector and a measurement equation of a corresponding subsystem under a special European group; resolving the state matrix, the measurement equation and the observation vector by using a Kalman filter to obtain navigation error estimation data; and performing feedback correction on the navigation error based on the navigation error estimation data.

Description

Technical Field

[0001] The present invention relates to the field of inertial navigation technology, and in particular to a launch system integrated navigation method of an inertial navigation / radar / satellite navigation triple combination based on Lie groups. Background Art

[0002] In related technologies, navigation systems need to be able to process information from multiple sources to ensure anti-interference capabilities in complex electromagnetic environments. Among existing navigation technology solutions, inertial navigation systems offer advantages such as comprehensive navigation information, good continuity, and high update frequency, but errors accumulate over time. Satellite navigation systems can provide highly accurate navigation positioning and timing, but are susceptible to external interference and have poor dynamic performance. Ground-based radar systems offer strong anti-interference capabilities, but their navigation information suffers from systematic errors, limited range, and requires datalink upload of position and velocity data. A single navigation system cannot meet the requirements of aerial self-alignment and integrated navigation in complex electromagnetic environments. In typical scenarios for aerial self-alignment and integrated navigation of artillery shells, at least two information sources, satellite and radar, are required to assist the inertial navigation system. Radar information ensures available navigation information within seconds to over ten seconds after launch, enabling rapid alignment if the satellite guidance system is interfered with or signal acquisition is incomplete. Satellite guidance, when available, ensures mid-flight and terminal impact accuracy, resulting in a navigation system with multi-dimensional information and higher navigation accuracy.

[0003] Traditional integrated navigation methods utilize filtering models built through a linearization process based on the assumption of small perturbation errors. However, these methods are limited by two factors under large misalignment conditions: first, the linearization assumptions no longer hold, and second, the origin of the linearized expansion is no longer accurately located. These two factors can cause the integrated navigation model to deviate from the actual model, leading to non-convergence at large misalignment angles and failure to meet the requirements for in-flight alignment and navigation.

[0004] It should be noted that the information disclosed in the above background technology section is only used to enhance the understanding of the background of the present invention, and therefore may include information that does not constitute prior art known to ordinary technicians in this field. Summary of the Invention

[0005] The present invention provides a transmitting system integrated navigation method for an inertial navigation / radar / satellite navigation triple combination based on a Lie group, a storage medium, a computer program product, and an electronic device. The method comprehensively utilizes inertial navigation, radar, and satellite information, establishes a state space model based on a special Euclidean group, achieves the effect of comprehensive and optimal utilization of multi-source information, and achieves rapid convergence under large misalignment angles, thereby overcoming the defects existing in the prior art to a certain extent.

[0006] Other features and advantages of the present invention will become apparent from the following detailed description, or may be learned in part by practice of the present invention.

[0007] According to a first aspect of the present invention, a launch system integrated navigation method for an inertial navigation / radar / satellite navigation triple combination based on Lie groups is provided, the method comprising:

[0008] Responding to navigation control instructions, configuring the aircraft's navigation system to enter an integrated navigation state;

[0009] A special Euclidean group is constructed based on posture features, velocity features and position features, and a motion description model based on the special Euclidean group description is established;

[0010] Obtain the navigation information of the inertial navigation system, solve the navigation information using the error model under the special Euclidean group, and obtain the state matrix;

[0011] Acquiring current measurement information; wherein the measurement information includes: satellite navigation subsystem measurement information and / or radar navigation subsystem measurement information;

[0012] Based on the measurement information, establishing an observation vector and a measurement equation of the corresponding subsystem under a special Euclidean group;

[0013] Solving the state matrix, measurement equation, and observation vector using a Kalman filter to obtain navigation error estimation data;

[0014] Feedback correction is performed on the navigation error based on the navigation error estimation data.

[0015] In some exemplary embodiments, the method further comprises:

[0016] Define the Earth-centered Earth-fixed coordinate system, launch coordinate system, carrier coordinate system, and rotation matrix corresponding to the aircraft;

[0017] Combine the rotation matrix, the transformation relationship between the launch coordinate system and the carrier coordinate system to determine the transformation matrix from the launch coordinate system to the carrier coordinate system

[0018] In some exemplary embodiments,

[0019] A special Euclidean group is constructed based on posture features, velocity features, and position features, and a motion description model based on the special Euclidean group is established, including:

[0020] Transformation matrix from the launch coordinate system to the carrier coordinate system Define a special orthogonal group SO(3); Based on the characteristic orthogonal group, construct a special Euclidean group SE(3);

[0021] The velocity vector and the position vector are expanded in the special Euclidean group to obtain a dual straight space isometric group SE2 (3) for representing the expanded posture, which is configured as a motion description model based on the special Euclidean group description.

[0022] In some exemplary embodiments, the method further comprises:

[0023] Incorporate posture features, velocity features, and position features into the Lie group to obtain the Lie group state quantity;

[0024] The Lie group state quantities are inverted and combined with the Lie group state quantities to construct a launch system strapdown inertial navigation error model based on Lie group as an error model under a special Euclidean group; wherein the launch system strapdown inertial navigation error model includes attitude, velocity and position errors.

[0025] In some exemplary embodiments, the combined navigation state vector based on the left error model includes:

[0026]

[0027] Among them, φ l represents the carrier attitude error, represents the carrier velocity error, represents the carrier position error, ε b Indicates the zero bias of the inertial device gyroscope, It represents the zero bias of the inertial device accelerometer, δA represents the radar azimuth measurement error, δE represents the radar elevation and low angle measurement error, and δR represents the radar radial distance measurement error.

[0028] Time updates are performed according to the state equation.

[0029] In some exemplary embodiments, the method further comprises:

[0030] Based on the inertial navigation subsystem and satellite navigation subsystem, the corresponding inertial / satellite integrated navigation measurement equations are defined, including: Among them, H vp is the measurement matrix, X is the combined navigation state vector based on the left error model, V vp is the measurement noise vector;

[0031] Based on the inertial navigation subsystem and radar navigation subsystem, the corresponding inertial navigation / radar integrated navigation measurement equations are defined, including: Among them, H rd is the measurement matrix, V rd is the measurement noise vector.

[0032] For satellites or radars, updates are measured.

[0033]

[0034] In some exemplary embodiments, the method further comprises:

[0035] After completing the current measurement update, determining whether there is measurement information corresponding to the satellite navigation subsystem and / or radar navigation subsystem to be processed;

[0036] If there is measurement information to be processed, continue with the measurement update; or

[0037] If there is no measurement information to be processed, the inertial navigation state is corrected.

[0038] According to a second aspect of the present invention, there is provided a computer program product having a computer program stored thereon, which, when executed by a processor, implements the above-mentioned Lie group-based inertial navigation / radar / satellite navigation three-combination launch system integrated navigation method.

[0039] According to a third aspect of the present invention, there is provided a storage medium having a computer program stored thereon, which, when executed by a processor, implements the above-mentioned Lie group-based inertial navigation / radar / satellite navigation three-combination launch system integrated navigation method.

[0040] According to a fourth aspect of the present invention, there is provided an electronic device, comprising:

[0041] a processor; and a memory for storing executable instructions for the processor;

[0042] Among them, the processor is configured to implement the above-mentioned Lie group-based inertial navigation / radar / satellite navigation three-combination launch system integrated navigation method by executing the executable instructions.

[0043] The Lie group-based integrated navigation method for a three-way inertial navigation / radar / satellite navigation system provided by an embodiment of the present invention establishes a state-space model based on a special Euclidean group, pre-constructing navigation measurement equations for the inertial navigation and radar combination, as well as for the inertial navigation and satellite navigation combination. This method allows for integrated navigation to be performed using the measurement information of the corresponding subsystems according to the specific type of measurement information corresponding to the current cycle, comprehensively utilizing navigation information from inertial navigation, satellites, and radar, achieving comprehensive and optimal utilization of multi-source information and improving navigation accuracy. This method significantly improves navigation convergence when processing navigation information from the inertial navigation system with large misalignment errors.

[0044] It is to be understood that the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] The accompanying drawings are incorporated into and constitute a part of this specification, illustrate embodiments consistent with the present invention, and together with the description, serve to explain the principles of the present invention. Obviously, the drawings described below are only some embodiments of the present invention, and it is clear that those skilled in the art can derive other drawings based on these drawings without inventive effort.

[0046] Figure 1 A schematic diagram schematically illustrates an exemplary embodiment of the present invention, a Lie group-based integrated navigation method for a transmitting system of an inertial navigation / radar / satellite navigation triple combination;

[0047] Figure 2 A schematic diagram schematically illustrating a relationship between a transmitting coordinate system and an Earth-centered Earth-fixed coordinate system according to an exemplary embodiment of the present invention;

[0048] Figure 3 A schematic diagram schematically illustrates data processing of an integrated navigation system according to an exemplary embodiment of the present invention;

[0049] Figure 4 A schematic diagram schematically illustrates a flow chart of an inertial navigation / radar / satellite navigation integrated navigation algorithm based on a Lie group error model according to an exemplary embodiment of the present invention;

[0050] Figure 5 The figure schematically shows the composition of an electronic device in an exemplary embodiment of the present invention. DETAILED DESCRIPTION

[0051] Example embodiments will now be described more fully with reference to the accompanying drawings. However, example embodiments can be embodied in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete and will fully convey the concepts of the example embodiments to those skilled in the art. The described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments.

[0052] In addition, the accompanying drawings are merely schematic illustrations of the present invention and are not necessarily drawn to scale. Identical reference numerals in the figures denote identical or similar parts, and thus repetitive descriptions thereof will be omitted. Some of the blocks shown in the accompanying drawings are functional entities that do not necessarily correspond to physically or logically separate entities. These functional entities may be implemented in software, in one or more hardware modules or integrated circuits, or in different networks and / or processor devices and / or microcontroller devices.

[0053] In view of the shortcomings and deficiencies of the existing technology, this example embodiment provides a launch system integrated navigation method based on Lie group inertial navigation / radar / satellite navigation three-combination. Figure 1As shown, the launch system integrated navigation method of the Lie group-based inertial navigation / radar / satellite navigation three-combination may specifically include the following steps:

[0054] Step S11, responding to the navigation control instruction, configuring the aircraft's navigation system to enter an integrated navigation state;

[0055] Step S12, constructing a special Euclidean group based on the posture features, velocity features, and position features, and establishing a motion description model based on the special Euclidean group description;

[0056] Step S13, obtaining navigation information of the inertial navigation system, solving the navigation information using an error model under a special Euclidean group, and obtaining a state matrix;

[0057] Step S14, obtaining current measurement information; wherein the measurement information includes: satellite navigation subsystem measurement information and / or radar navigation subsystem measurement information;

[0058] Step S15: establishing an observation vector and a measurement equation for the corresponding subsystem under a special Euclidean group based on the measurement information;

[0059] Step S16, using a Kalman filter to solve the state matrix, measurement equation, and observation vector to obtain navigation error estimation data;

[0060] Step S17: performing feedback correction on the navigation error based on the navigation error estimation data.

[0061] Below, each step of the Lie group-based inertial navigation / radar / satellite navigation triple combination transmitting system integrated navigation method in this example implementation will be described in more detail with reference to the accompanying drawings and embodiments.

[0062] In step S11, in response to the navigation control instruction, the navigation system of the aircraft is configured to enter the integrated navigation state.

[0063] For example, the aforementioned aircraft may be a drone, airplane, or missile. The aircraft may be pre-installed with an inertial navigation system, a satellite navigation system, and a radar navigation system. During flight or before takeoff, the aircraft can switch to an integrated navigation state and execute an integrated navigation mode based on navigation control commands issued by the user. In this integrated navigation mode, the aircraft can optimize navigation information based on the navigation parameters of the inertial navigation system, satellite navigation system, and radar navigation system.

[0064] Alternatively, in some exemplary embodiments, the accuracy of the inertial navigation system's navigation parameters may be evaluated in real time during flight. When the error in the inertial navigation system's navigation parameters exceeds a preset threshold, a navigation control instruction is automatically generated, causing the aircraft's navigation system to enter an integrated navigation state. For example, a navigation control instruction may be triggered when the deviation angle between the inertial navigation system's navigation coordinate system and the true navigation coordinate system exceeds a preset threshold.

[0065] In step S12, a special Euclidean group is constructed based on the posture features, velocity features, and position features, and a motion description model based on the special Euclidean group description is established.

[0066] Exemplarily, the method further includes:

[0067] Define the Earth-centered Earth-fixed coordinate system, launch coordinate system, carrier coordinate system, and rotation matrix corresponding to the aircraft;

[0068] Combine the rotation matrix, the transformation relationship between the launch coordinate system and the carrier coordinate system to determine the transformation matrix from the launch coordinate system to the carrier coordinate system

[0069] Specifically, we can first establish a kinematic model for the aircraft and agree on a related coordinate system. The aircraft-related coordinate system may include:

[0070] (1) Earth-centered Earth-fixed coordinate system (e system)

[0071] The Earth-centered, Earth-fixed coordinate system, also known as the e-system, has its origin at the center of the Earth, the x-axis in the equatorial plane pointing to the prime meridian, the z-axis for the Earth's rotation axis pointing to the North Pole, and the y-axis in the equatorial plane forming a right-handed rectangular coordinate system with the x-axis and the z-axis.

[0072] (2) Emission coordinate system (g system)

[0073] The launch coordinate system, that is, the g system, the coordinate origin O g The launch point is the launch point. The x-axis points to the launch aiming direction in the horizontal plane of the launch point. The y-axis is perpendicular to the horizontal plane of the launch point and points upward. The z-axis, x-axis, and y-axis form a right-handed rectangular coordinate system. The launch system is fixed to the earth. The geographic latitude B0, longitude λ0, altitude h0, and launch azimuth A0 of the launch point determine the relationship between the launch coordinate system and the earth, as shown in the following example: Figure 2 shown.

[0074] (3) Carrier coordinate system

[0075] The vehicle coordinate system for the aircraft / carrier, or b-frame, has its origin at the center of mass of the aircraft. For example, if the aircraft is a drone or a missile, the x-axis coincides with the longitudinal axis of the missile, the y-axis lies in the longitudinal plane of symmetry and is perpendicular to the x-axis, and the z-axis forms a right-handed rectangular coordinate system with the x- and y-axes.

[0076] (4) Rotation Matrix

[0077] The rotation matrices for rotating α around the x-axis, y-axis, and z-axis are:

[0078]

[0079] (5) Launch coordinate system and carrier coordinate system

[0080] Transformation matrix from launch coordinate system to carrier coordinate system Expressed as:

[0081]

[0082] Among them, R x (γ) represents, R y (ψ) represents, express,.

[0083] Transformation matrix from carrier coordinate system to launch coordinate system Expressed as:

[0084]

[0085] In step S12, a special Euclidean group is constructed based on the posture features, velocity features, and position features, and a motion description model based on the special Euclidean group description is established.

[0086] Exemplarily, the above step S12 may include:

[0087] Transformation matrix from the launch coordinate system to the carrier coordinate system Define a special orthogonal group So(3); Based on the characteristic orthogonal group, construct a special Euclidean group SE(3);

[0088] The velocity vector and the position vector are expanded in the special Euclidean group to obtain a dual straight space isometric group SE2 (3) for representing the expanded posture, which is configured as a motion description model based on the special Euclidean group description.

[0089] Exemplarily, the method further includes:

[0090] Incorporate posture features, velocity features, and position features into the Lie group to obtain the Lie group state quantity;

[0091] The Lie group state quantities are inverted and combined with the Lie group state quantities to construct a launch system strapdown inertial navigation error model based on Lie group as an error model under a special Euclidean group; wherein the launch system strapdown inertial navigation error model includes attitude, velocity and position errors.

[0092] For example, the navigation posture, velocity and position may be incorporated into a special Euclidean group, and a motion description model based on the special Euclidean group description may be established.

[0093] Specifically, a Lie group is an algebraic structure consisting of sets and operations, where the group operations satisfy the property of continuous differentiability. Lie groups are mathematical objects that possess both the structure of an algebraic group and the properties of a differential manifold. The combination of their smooth manifold properties and continuous group operations gives Lie groups a unique advantage in solving differential equations.

[0094] The direction cosine matrix (that is, the above transformation matrix ) is defined as the special orthogonal group SO(3), and the formula is expressed as:

[0095]

[0096] The navigation attitude, velocity and position are defined as the special Euclidean group SE(3), which can be expressed as follows:

[0097]

[0098] in, Represents a 3×3 dimensional space; represents a group variable; V g Indicates the velocity of the launch system.

[0099] The carrier motion can be regarded as a coordinate system that is transformed into another coordinate system through rotation and translation. The translation can be represented by a three-dimensional vector, and the rotation can be represented by a 3×3 matrix R. The attitude matrix R satisfies the following formula

[0100]

[0101] Among them, I 3×3 is a 3×3 unit matrix, Represents a 3×3 dimensional space.

[0102] The velocity vector V and position vector P are extended to SE(3) to represent the double straight space isometric group SE2(3) of the extended posture as the motion description model. The formula is expressed as

[0103]

[0104] Among them, 0 1×3 Represents a 1×3 order zero matrix.

[0105] For SE2(3), its reciprocal is given by:

[0106]

[0107] Therefore, the reciprocal of SE2(3) also belongs to SE2(3).

[0108] Exemplarily, the method further includes: incorporating posture features, velocity features, and position features into a Lie group to obtain a Lie group state quantity;

[0109] The Lie group state quantities are inverted and combined with the Lie group state quantities to construct a launch system strapdown inertial navigation error model based on Lie group as an error model under a special Euclidean group; wherein the launch system strapdown inertial navigation error model includes attitude, velocity and position errors.

[0110] Specifically, a strapdown inertial navigation error model based on Lie group can be constructed to convert the attitude speed and position P g Incorporating it into the Lie group can form the Lie group state quantity, and the formula can be expressed as:

[0111]

[0112] For the above formula (9), the Lie group state quantity χ is inverted to obtain

[0113]

[0114] The error model of the launch system strapdown inertial navigation based on Lie group is defined as:

[0115]

[0116] in, and is the navigation parameter with error obtained by the inertial navigation system, namely attitude, velocity, and position parameters; τ a , τ v and τ p are the attitude, velocity and position errors in the strapdown inertial navigation error model of the launch system.

[0117] Let φ be the misalignment angle error, and the antisymmetric matrix of φ is ψ, that is, ψ = [φ × ]. When φ is at its minimum, according to the relationship between Lie groups and Lie algebras, the attitude, velocity, and position errors can be expressed as:

[0118]

[0119] in, represents the auxiliary velocity vector in the launch coordinate system,

[0120] In step S13, the navigation information of the inertial navigation system is obtained, and the navigation information is solved using an error model under a special Euclidean group to obtain a state matrix.

[0121] For example, the attitude, velocity and position error differential equations based on the Lie group error model are:

[0122]

[0123] in, δf b Respectively expressed as:

[0124]

[0125] Among them, w a represents the added measurement noise, w g represents the gyro measurement noise.

[0126] The combined navigation state vector based on the left error model can be expressed as:

[0127]

[0128] Among them, φ l represents the carrier attitude error under the inertial navigation system, represents the carrier velocity error, represents the carrier position error, ε b Indicates the zero bias of the inertial device gyroscope, It represents the zero bias of the inertial device accelerometer, δA represents the radar azimuth measurement error, δE represents the radar elevation and low angle measurement error, and δR represents the radar radial distance measurement error.

[0129] Based on the above equations (15) and (16), the combined navigation state equation of the left error model can be obtained as follows:

[0130]

[0131] Among them, F l is the state transfer matrix, G l is the noise driving matrix, W l is the process noise vector.

[0132] The above formula (18) can be expressed as:

[0133]

[0134] In step S14, current measurement information is acquired; wherein the measurement information includes: satellite navigation subsystem measurement information and / or radar navigation subsystem measurement information.

[0135] For example, the aircraft may obtain measurement information for the current period. The current measurement information may include satellite navigation subsystem measurement information, radar navigation subsystem measurement information, or both. The aircraft may identify the type of the current measurement information.

[0136] In step S15, based on the measurement information, an observation vector and a measurement equation of the corresponding subsystem under a special Euclidean group are established.

[0137] For example, the inertial / satellite integrated navigation measurement equation is:

[0138] Z vp =H vp X+V vp (20)

[0139] Among them, H vp As shown in formula (23); V vp is the measurement noise vector of the measurement equation.

[0140] The speed and position measurement vector is:

[0141]

[0142] Among them, V I is the auxiliary velocity vector output by inertial navigation, P I V is the position vector output by inertial navigation; S is the auxiliary velocity vector output by satellite navigation, P S It is the auxiliary position vector output by satellite navigation.

[0143] According to the above formula (13) and formula (14), we can get:

[0144]

[0145] Then the corresponding measurement matrix H vp The formula can be expressed as:

[0146]

[0147] For example, the inertial navigation / radar integrated navigation measurement equation can be expressed as:

[0148]

[0149] Z rd =H rd X+V rd (25)

[0150] The corresponding measurement matrix Hrd The expression is:

[0151]

[0152] in,

[0153]

[0154] Among them, the current measurement matrix H rd The right side of H1, H2, and H3 are all This is because the current combined navigation state quantity is redefined by the Lie group and satisfies the relationship shown in formula (27).

[0155]

[0156] Therefore, it is necessary to map the current measurement matrix to the currently defined state quantity through posture transformation.

[0157] In step S16, the state matrix, measurement equation, and observation vector are solved using a Kalman filter to obtain navigation error estimation data.

[0158] Exemplary, reference Figure 3 As shown, the observation vector and measurement equation of the current subsystem, as well as the state equation corresponding to the inertial navigation system, are input into the Kalman filter and solved, thereby obtaining the optimal estimate of the navigation error output by the Kalman filter.

[0159] For example, the velocity position measurement vector and measurement matrix corresponding to the above-mentioned inertial navigation / satellite navigation combination, i.e., the calculation results of the above-mentioned formula (24) and formula (26), are input into the Kalman filter. Alternatively, the radar measurement vector and measurement matrix H corresponding to the above-mentioned inertial navigation / radar combination are input into the Kalman filter. rd , that is, the calculation results of the above formulas (28) and (29) are input into the Kalman filter.

[0160] In step S17 , feedback correction is performed on the navigation error based on the navigation error estimation data.

[0161] For example, after obtaining the navigation error estimation data output by the Kalman filter, it can be fed back to the inertial navigation system to correct the navigation information of the inertial navigation system.

[0162] Exemplarily, the method further includes: after completing the current measurement update, determining whether there is measurement information corresponding to the satellite navigation subsystem and / or the radar navigation subsystem to be processed;

[0163] If there is measurement information to be processed, continue with the measurement update; or

[0164] If there is no measurement information to be processed, the inertial navigation state is corrected.

[0165] For example, the navigation system installed in the aircraft includes a GNSS receiving module, a gyroscope, an accelerometer, a radar receiving module, and a navigation calculation module. The navigation calculation module is used to collect data from the GNSS, gyroscope, accelerometer, and radar, and perform inertial navigation solution and Kalman filter calculation. Figure 4 As shown, the following steps may be specifically included:

[0166] S1. The navigation system enters the combined navigation state.

[0167] S2. Establish a special Euclidean group description error model.

[0168] Specifically, the navigation posture, velocity and position are incorporated into the special Euclidean group, and a motion description model based on the special Euclidean group description is established.

[0169] S3. Collect inertial navigation information and solve the state matrix.

[0170] Specifically, the navigation information of the inertial navigation system is obtained, substituted into the error model under the special Euclidean group, and the time update of the state matrix is ​​solved.

[0171] The corresponding Kalman filter formula includes:

[0172]

[0173] Update the time according to the state equation. Let the state at time k (k = 0, 1, 2...; time interval is T) be X k , the variance of the state is recorded as P k , can be obtained from X and P (denoted as X) at time k-1 k-1 and P k-1 ) for X and P at time k (denoted as X k / k-1 and P k / k-1 ) to make predictions.

[0174] Φ k / k-1 =I+F l T

[0175] Where I is the identity matrix, and its dimension is the same as F.

[0176] S4. Collect observations and construct measurement equations.

[0177] Specifically, the system acquires measurement information from satellite navigation and / or radar, determines its validity, captures information from any subsystem of the satellite navigation and radar, and establishes the observation vector and measurement equation for the corresponding subsystem under a special Euclidean group.

[0178] S5. Measurement update.

[0179] Specifically, the state equation, measurement equation and observation are substituted into the Kalman filter to solve the optimal estimation of the navigation error.

[0180] For satellites or radars, the measurement matrix H vp or H rd Substitute H k , complete the measurement update according to the following formula.

[0181]

[0182] S6. Navigation error feedback correction.

[0183] Specifically, after completing a measurement update, check whether there is any satellite navigation and radar information to be processed. If so, continue to perform measurement update; if not, correct the inertial navigation state.

[0184] The method provided by the embodiment of the present invention can fuse the navigation information of inertial navigation, satellite and radar, and can significantly improve the navigation convergence under the condition that there is a large misalignment angle error in processing the navigation information of the inertial navigation system. It should be noted that the above-mentioned figures are only schematic illustrations of the processing included in the method according to the exemplary embodiment of the present invention, and are not for limiting purposes. It is easy to understand that the processing shown in the above-mentioned figures does not indicate or limit the time sequence of these processing. In addition, it is also easy to understand that these processing can be performed synchronously or asynchronously, for example, in multiple modules.

[0185] It should be noted that, although several modules or units of the device for action execution are mentioned in the above detailed description, this division is not mandatory. In fact, according to an embodiment of the present invention, the features and functions of two or more modules or units described above can be concretized in one module or unit. Conversely, the features and functions of one module or unit described above can be further divided into multiple modules or units to be concretized.

[0186] Figure 5 A schematic diagram of an electronic device suitable for implementing an embodiment of the present invention is shown.

[0187] It should be noted that Figure 5 The electronic device 1000 shown is only an example and should not limit the functions and scope of use of the embodiments of the present invention.

[0188] For example, the electronic device may be an intelligent electronic device installed on an aircraft and used to navigate the aircraft.

[0189] like Figure 5As shown, electronic device 1000 includes a central processing unit (CPU) 1001, which can perform various appropriate actions and processes according to the program stored in read-only memory (ROM) 1002 or the program loaded from storage portion 1008 into random access memory (RAM) 1003. Various programs and data required for system operation are also stored in RAM 1003. CPU 1001, ROM 1002 and RAM 1003 are connected to each other via bus 1004. Input / output (I / O) interface 1005 is also connected to bus 1004.

[0190] The following components are connected to the I / O interface 1005: an input section 1006 including a keyboard, a mouse, and the like; an output section 1007 including devices such as a cathode ray tube (CRT), a liquid crystal display (LCD), and a speaker; a storage section 1008 including a hard disk and the like; and a communication section 1009 including a network interface card such as a LAN (Local Area Network) card or a modem. The communication section 1009 performs communication processing via a network such as the Internet. A drive 1010 is also connected to the I / O interface 1005 as needed. Removable media 1011, such as a magnetic disk, an optical disk, a magneto-optical disk, or a semiconductor memory, is installed in the drive 1010 as needed, so that computer programs read therefrom can be installed into the storage section 1008 as needed.

[0191] In particular, according to an embodiment of the present invention, the process described below with reference to the flowchart can be implemented as a computer software program. For example, an embodiment of the present invention includes a computer program product that includes a computer program carried on a storage medium, the computer program containing program code for executing the method shown in the flowchart. In such an embodiment, the computer program can be downloaded and installed from a network via the communication section 1009 and / or installed from a removable medium 1011. When the computer program is executed by the central processing unit (CPU) 1001, the various functions defined in the system of the present application are performed.

[0192] It should be noted that the storage medium shown in the embodiments of the present invention can be a computer-readable signal medium or a computer-readable storage medium or any combination of the above. The computer-readable storage medium can be, for example, but not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, device or component, or any combination of the above. More specific examples of computer-readable storage media can include, but are not limited to: an electrical connection with one or more wires, a portable computer disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM), a flash memory, an optical fiber, a portable compact disc read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the above. In the present invention, a computer-readable storage medium can be any tangible medium that contains or stores a program that can be used by or in conjunction with an instruction execution system, device or device. In the present invention, a computer-readable signal medium can include a data signal propagated in baseband or as part of a carrier wave, which carries computer-readable program code. Such propagated data signals may take various forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination thereof. A computer-readable signal medium may also be any storage medium other than a computer-readable storage medium that can transmit, propagate, or transport a program for use by or in connection with an instruction execution system, apparatus, or device. Program code contained on the storage medium may be transmitted using any suitable medium, including but not limited to wireless, wired, or any suitable combination thereof.

[0193] The flowcharts and block diagrams in the accompanying drawings illustrate the possible implementation architecture, functions and operations of the systems, methods and computer program products according to various embodiments of the present invention. In this regard, each box in the flowchart or block diagram can represent a module, program segment, or a part of code, and the above-mentioned module, program segment, or a part of code contains one or more executable instructions for implementing the specified logical function. It should also be noted that in some alternative implementations, the functions marked in the box can also occur in an order different from that marked in the accompanying drawings. For example, two boxes represented in succession can actually be executed substantially in parallel, and they can sometimes be executed in the opposite order, depending on the functions involved. It should also be noted that each box in the block diagram or flowchart, and the combination of boxes in the block diagram or flowchart, can be implemented with a dedicated hardware-based system that performs the specified function or operation, or can be implemented with a combination of dedicated hardware and computer instructions.

[0194] The units involved in the embodiments of the present invention may be implemented in software or hardware, and the units described may also be provided in a processor. In some cases, the names of these units do not limit the units themselves.

[0195] It should be noted that, as another aspect, the present application also provides a storage medium, which can be included in an electronic device; or it can exist independently without being installed in the electronic device. The above storage medium carries one or more programs, and when the above one or more programs are executed by an electronic device, the electronic device implements the method described in the following embodiments. For example, the electronic device can implement the following Figure 1 The individual steps of the method are shown.

[0196] In one embodiment, the present application provides a computer program product, including a computer program, which implements the steps in the above-mentioned method embodiments when executed by a processor.

[0197] Furthermore, the above-described figures are merely illustrative of the processes included in the method according to exemplary embodiments of the present invention and are not intended to be limiting. It is readily understood that the processes illustrated in the above-described figures do not indicate or limit the temporal order of these processes. Furthermore, it is readily understood that these processes may be executed synchronously or asynchronously, for example, in multiple modules.

[0198] Other embodiments of the present invention will readily occur to those skilled in the art after considering the specification and practicing the invention herein. This application is intended to cover any variations, uses, or adaptations of the present invention that follow the general principles of the invention and include common knowledge or customary techniques in the art not disclosed herein. The description and examples are to be considered as exemplary only, with the true scope and spirit of the invention being indicated by the claims.

[0199] It should be understood that the present invention is not limited to the exact construction described above and shown in the drawings, and that various modifications and changes may be made without departing from the scope thereof, which is limited only by the appended claims.

Claims

1. A Lie group-based integrated navigation method for a transmitting system of an inertial navigation / radar / satellite navigation triple combination, characterized in that: The method comprises: Responding to navigation control instructions, configuring the aircraft's navigation system to enter an integrated navigation state; A special Euclidean group is constructed based on posture features, velocity features and position features, and a motion description model based on the special Euclidean group description is established; Obtain the navigation information of the inertial navigation system, solve the navigation information using the error model under the special Euclidean group, and obtain the state matrix; Acquiring current measurement information; wherein the measurement information includes: satellite navigation subsystem measurement information and / or radar navigation subsystem measurement information; Based on the measurement information, establishing an observation vector and a measurement equation of the corresponding subsystem under a special Euclidean group; Solving the state matrix, measurement equation, and observation vector using a Kalman filter to obtain navigation error estimation data; Feedback correction is performed on the navigation error based on the navigation error estimation data.

2. The method according to claim 1, characterized in that The method further comprises: Define the Earth-centered Earth-fixed coordinate system, launch coordinate system, carrier coordinate system, and rotation matrix corresponding to the aircraft; Combine the rotation matrix, the transformation relationship between the launch coordinate system and the carrier coordinate system to determine the transformation matrix from the launch coordinate system to the carrier coordinate system 3. The method according to claim 2, characterized in that A special Euclidean group is constructed based on posture features, velocity features, and position features, and a motion description model based on the special Euclidean group is established, including: Transformation matrix from the launch coordinate system to the carrier coordinate system Define the special orthogonal group So(3); Based on the characteristic orthogonal group, construct the special Euclidean group SE(3); The velocity vector and the position vector are expanded in the special Euclidean group to obtain a dual straight space isometric group SE2 (3) for representing the expanded posture, which is configured as a motion description model based on the special Euclidean group description.

4. The method according to claim 3, characterized in that The method further comprises: Incorporate posture features, velocity features, and position features into the Lie group to obtain the Lie group state quantity; The Lie group state quantities are inverted and combined with the Lie group state quantities to construct a launch system strapdown inertial navigation error model based on Lie group as an error model under a special Euclidean group; wherein the launch system strapdown inertial navigation error model includes attitude, velocity and position errors.

5. The method according to claim 1, wherein The method further comprises: Based on the inertial navigation subsystem and satellite navigation subsystem, the corresponding inertial / satellite integrated navigation measurement equations are defined, including: Among them, H vp is the measurement matrix, X is the combined navigation state vector based on the left error model, V vp is the measurement noise vector; Based on the inertial navigation subsystem and radar navigation subsystem, the corresponding inertial navigation / radar integrated navigation measurement equations are defined, including: Z rd =H rd X+V rd Among them, H rd is the measurement matrix, V rd is the measurement noise vector.

6. The method according to claim 1, characterized in that The combined navigation state vector based on the left error model includes: Among them, φ l represents the carrier attitude error, represents the carrier velocity error, represents the carrier position error, ε b Indicates the zero bias of the inertial device gyroscope, It represents the zero bias of the inertial device accelerometer, δA represents the radar azimuth measurement error, δE represents the radar elevation and low angle measurement error, and δR represents the radar radial distance measurement error.

7. The method according to claim 1, characterized in that The method further comprises: After completing the current measurement update, determining whether there is measurement information corresponding to the satellite navigation subsystem and / or radar navigation subsystem to be processed; If there is measurement information to be processed, continue with the measurement update; or If there is no measurement information to be processed, the inertial navigation state is corrected.

8. A storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the transmitting system integrated navigation method of the Lie group-based inertial navigation / radar / satellite navigation three-combination is implemented as described in any one of claims 1 to 7.

9. A computer program product comprising a computer program, characterized in that When the computer program is executed by a processor, the transmitting system integrated navigation method of the inertial navigation / radar / satellite navigation triple combination based on Lie group according to any one of claims 1 to 7 is implemented.

10. An electronic device, characterized in that: include: processor; as well as a memory for storing executable instructions of the processor; The processor is configured to execute the Lie group-based inertial navigation / radar / satellite navigation triple combination launch system integrated navigation method according to any one of claims 1 to 7 by executing the executable instructions.

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