High-precision alignment method for airborne inertial navigation

Abstract

The application relates to a high-precision alignment method of airborne inertial navigation, which comprises the following steps: step 1, selecting a coarse alignment mode according to an initial alignment scene selected by a user; step 2, performing inertial navigation coarse alignment according to the coarse alignment mode selected in step 1; step 3, after the inertial navigation coarse alignment in step 2 is completed, obtaining a rough initial attitude angle and a position cosine matrix, and then performing fine alignment; and step 4, when the airplane maneuvering is completed, the inertial navigation completes the fine alignment process by using a Kalman filtering algorithm, the attitude angle error estimated by the Kalman filtering is used to correct the navigation solution attitude angle of the strapdown inertial navigation itself, after the fine alignment is completed, the inertial navigation enters a navigation state, and the corrected attitude and speed information are output, and then the initial alignment process is completed. The application can realize high-precision alignment and does not need the assistance of high-precision master inertial navigation or high-precision satellite navigation RTK information.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of airborne or missile-borne inertial navigation system, and relates to a high-precision alignment method, in particular to a high-precision alignment method of airborne inertial navigation. BACKGROUND

[0002] Due to the requirements of miniaturization and low power consumption of a strapdown inertial navigation system in a high-maneuvering environment such as an airborne or missile-borne environment, the strapdown inertial navigation system equipped in the environment is usually composed of low-precision miniaturized gyroscopes and accelerometers, and therefore the initial alignment cannot meet the requirements of high-precision surveying and mapping and attitude measurement in the military field due to poor alignment precision without the assistance of main inertial navigation auxiliary alignment.

[0003] Through retrieval, no existing technical patent document similar to the application is found. SUMMARY

[0004] The application aims to overcome the problem of insufficient theoretical support system based on experience in the prior art, and proposes a high-precision alignment method of airborne inertial navigation, which can realize high-precision alignment without the assistance of high-precision main inertial navigation or high-precision satellite navigation RTK information, and the initial alignment precision is improved by one order of magnitude compared with the traditional shaking and static initial alignment environment.

[0005] The application solves the practical problems by adopting the following technical scheme:

[0006] A high-precision alignment method of airborne inertial navigation comprises the following steps:

[0007] Step 1: selecting a coarse alignment mode according to the initial alignment scene selected by a user;

[0008] Step 2: performing inertial navigation coarse alignment according to the coarse alignment mode selected in step 1;

[0009] Step 3: after the inertial navigation coarse alignment in step 2 is completed, a rough initial attitude angle and a position cosine matrix are obtained, and then fine alignment is performed; at the same time, the aircraft is flown in a set maneuvering mode, and if the ground alignment is selected in step 1, the aircraft only needs to start normal gliding take-off on the ground and does not need to be specially maneuvered;

[0010] Step 4: after the aircraft maneuvering is completed, the inertial navigation is completed by using a Kalman filtering algorithm to complete the fine alignment process, the attitude angle error estimated by the Kalman filtering is used to correct the navigation solution attitude angle of the strapdown inertial navigation system, and after the fine alignment is completed, the inertial navigation system enters the navigation state to output the corrected attitude and speed information, thereby completing the initial alignment process.

[0011] Moreover, the specific method of step 1 is:

[0012] If ground alignment is selected, the inertial navigation dynamic coarse alignment algorithm is converted to the solid coordinate system coarse alignment algorithm; if air alignment is selected, the Beidou track angle information is input to perform the coarse alignment of the track angle;

[0013] Moreover, the specific method of the step 2 is:

[0014] (1) If ground alignment is selected, the inertial navigation dynamic coarse alignment algorithm is converted to the solid coordinate system coarse alignment algorithm:

[0015] First, the azimuth cosine matrix is decomposed as:

[0016]

[0017] a) find

[0018]

[0019] b) find

[0020] Let The corresponding attitude quaternion is denoted as Convert to the attitude quaternion Form:

[0021]

[0022] When k = 0, that is, at the start time of the coarse alignment The b system coincides with the system, Therefore

[0023]

[0024] When k = 0, 1, 2,..., the attitude quaternion at time t is calculated by using the quaternion updating algorithm k+1

[0025]

[0026] In the formula:

[0027]

[0028] Let:

[0029]

[0030]

[0031] In the formula:

[0032] Δθ = [Δθ x Δθ​y Delta theta z ] T Gyro output angle increment in sampling period, unit: rad. k t k+1 ]Sampling period

[0033] Normalization processing

[0034]

[0035] Get updated attitude quaternion Using and conversion relationship to get t k+1 time

[0036]

[0037] Finally,

[0038] c) Get

[0039] Coarse alignment start time, that is, k=0,

[0040] k=0,1,2,...,

[0041]

[0042] In the formula:

[0043]

[0044] Take t n =t k+1 and t m =t k+1-N time (N=10) two points Get

[0045]

[0046] Substitute the above and into to get the azimuth cosine matrix;

[0047] (2) If air alignment is selected, input by Beidou track angle information, and perform coarse alignment of satellite navigation track angle: the angle defined as H after synchronization of satellite navigation track angle is transmitted to the inertial navigation system, and the initial horizontal attitude angle is zero, then the initial attitude angle is obtained as:

[0048] att0=[0 0H]

[0049] The azimuth cosine matrix of the coarse alignment is obtained by the attitude angle rotation cosine matrix formula:

[0050]

[0051] Moreover, the specific method of step 4 is as follows:

[0052] The state variable of the Kalman filter algorithm in the fine alignment is selected as follows:

[0053]

[0054] The selected state variable includes the attitude angle The eastward and northward velocity error δv of the two navigation channels x , δv y The position coordinate point error δx, δy, δz of the two navigation channels, the gyro constant drift ε x , ε y , ε z The constant bias of the accelerometer measurement

[0055] The general form of the system observation equation is

[0056]

[0057] In which, H is the system observation matrix, is the observation noise vector of the system.

[0058] The position of the satellite guidance is taken as the external observation information of the system, and the measurement equation of the observation quantity is shown in the formula

[0059]

[0060] In which:

[0061] H v is the observation equation transfer matrix;

[0062] V is the observation noise vector.

[0063] In which, the subscript Ins represents the latitude, longitude and height information calculated by the inertial navigation system, and the subscript r represents the latitude, longitude and height information output by the satellite guidance. H v = [0 3×6 I 3×3 0 3×6 ], and V is a 3-dimensional velocity observation white noise;

[0064] At this time, the state equation of the Kalman filter is as follows:

[0065]

[0066] Wherein, the F, G state transition matrix and noise matrix adopt the general mathematical model of inertial navigation algorithm.

[0067] Advantages and beneficial effects of the present application:

[0068] The present application proposes a high-precision alignment method for airborne inertial navigation, which utilizes the high-mobility special application environment in the actual airborne environment, and designs an initial alignment scheme using the carrier mobility condition under the auxiliary condition of single-point satellite navigation information. Various mobility schemes are designed to complete the initial alignment according to the use conditions of airborne inertial navigation. Unlike the traditional initial alignment precision of strapdown inertial navigation, which is restricted by the equivalent eastward gyro drift, the airborne inertial navigation in the present application can achieve high-precision alignment by relying only on satellite navigation single-point positioning method through the process of mobility excitation, without the assistance of high-precision main inertial navigation or high-precision satellite navigation RTK information. The initial alignment precision is improved by an order of magnitude compared with the traditional shaking and static initial alignment environment. BRIEF DESCRIPTION OF DRAWINGS

[0069] Figure 1 The flow chart of high-precision fine alignment of the airborne inertial navigation of the present application;

[0070] Figure 2 The flow chart of coarse alignment of the present application;

[0071] Figure 3 The flow chart of high-precision initial alignment of the airborne inertial navigation of the present application;

[0072] Figure 4 The alignment error curve chart of the present application. DETAILED DESCRIPTION

[0073] The embodiments of the present application will be further described in detail below in combination with the drawings:

[0074] A high-precision alignment method for airborne inertial navigation, as shown in Figure 1 , includes the following steps:

[0075] Step 1: Selecting coarse alignment method according to user-selected initial alignment scene;

[0076] The specific method of step 1 is:

[0077] If ground alignment is selected, the inertial navigation coarse alignment is performed by the solid coordinate system coarse alignment algorithm; if air alignment is selected, the Beidou track angle information is inputted, and the satellite navigation track angle coarse alignment is performed;

[0078] Step 2: Performing inertial navigation coarse alignment according to the coarse alignment method selected in step 1;

[0079] In the embodiment, coarse alignment or initial heading angle transfer process is performed according to gyro acceleration information output by the IMU, and when the satellite navigation track angle or position and velocity information is utilized, inertial navigation and satellite navigation information synchronization compensation is required, the synchronization compensation is realized by combining software and hardware, and since the synchronization algorithm is a common means in the navigation field, the present application will not be described here.

[0080] As shown in Figure 3 , the specific method of step 2 is:

[0081] (1) If ground alignment is selected, the inertial navigation dynamic coarse alignment is performed by converting into the solidified coordinate system coarse alignment algorithm:

[0082] First, the azimuth cosine matrix is decomposed into

[0083]

[0084] d) solve

[0085]

[0086] e) solve

[0087] Let the corresponding attitude quaternion be Convert into attitude quaternion in the form:

[0088]

[0089] When k=0, that is, at the beginning of coarse alignment The b system coincides with the b system, So

[0090]

[0091] When k=0, 1, 2,..., the attitude quaternion at time t is calculated by using the quaternion update algorithm k+1

[0092]

[0093] In the formula:

[0094]

[0095] Let:

[0096]

[0097] ​

[0098] where:

[0099] Δθ = [Δθ x Δθ y Δθ z ] T — (t k , t k+1 ] the gyro output angular increment in radian in the sampling period.

[0100] Normalization

[0101]

[0102] get the updated attitude quaternion using and the conversion relationship t k+1 time

[0103]

[0104] Finally,

[0105] f) get

[0106] At the beginning of the coarse alignment, that is, when k = 0,

[0107] When k = 0, 1, 2,...,

[0108]

[0109] where:

[0110]

[0111] Take t n = t k+1 and t m = t k+1-N time (N = 10) two points get

[0112]

[0113] Substitute the above and into , get the azimuth cosine matrix;

[0114] (2) If the air alignment is selected, the coarse alignment of the satellite navigation track angle is performed by inputting the Beidou track angle information: the angle defined as H after the satellite navigation track angle is transmitted to the inertial navigation after synchronization, and the initial attitude angle is zero, so that the initial attitude angle can be obtained as:

[0115] att0 = [0 0H]

[0116] The azimuth cosine matrix of the coarse alignment is obtained by the azimuth cosine matrix formula after the attitude angle conversion:

[0117]

[0118] Step 3: After the coarse alignment of the inertial navigation in step 2 is completed, the rough initial attitude angle and the azimuth cosine matrix are obtained, and then the fine alignment is performed; at the same time, the aircraft flies according to the set flight maneuver, if the ground alignment is selected in step 1, the aircraft only needs to start the normal gliding take-off on the ground, and does not need to be specially maneuvered;

[0119] The process of the aircraft flying according to the set flight maneuver in step 3 is shown in Table 1:

[0120] Table 1: Aircraft maneuver requirements during alignment

[0121]

[0122] The aircraft can select one or a combination of multiple maneuver modes according to the maneuvers listed in Table 1 to complete the requirements for the maneuver in the fine alignment process.

[0123] Step 4: After the aircraft maneuver is completed, the Kalman filter algorithm is used to complete the fine alignment process, and the attitude angle error estimated by the Kalman filter is used to correct the navigation solution attitude angle of the strapdown inertial navigation itself, after the fine alignment is completed, the inertial navigation enters the navigation state, and outputs the corrected attitude and speed information, thereby completing the initial alignment process;

[0124] The specific method of step 4 is:

[0125] The state variables of the Kalman filter algorithm in the fine alignment are as follows:

[0126]

[0127] The selected state variables include attitude angles Eastward and northward speed errors δv x , δv y of the two navigation channels x Position coordinate point errors δx, δy, δz of the two navigation channels y Gyro constant drifts ε z Accelerometer measurement constant bias

[0128] The general form of the system observation equation is

[0129]

[0130] Wherein, H is the system observation matrix, is the observation noise vector of the system.

[0131] The system takes the position of the satellite as the external observation information, and the measurement equation of the observation quantity is shown in the formula

[0132]

[0133] In the formula,

[0134] H v is the observation equation transfer matrix;

[0135] V is the observation noise vector.

[0136] In the formula, the subscript Ins represents the latitude, longitude and height information calculated and output by the inertial navigation system, and the subscript r represents the latitude, longitude and height information output by the satellite. H v = [0 3×6 I 3×3 0 3×6 ], and V is a 3-dimensional velocity observation white noise

[0137] At this time, the state equation of the Kalman filter is:

[0138]

[0139] Wherein, F and G are the state transition matrix and noise matrix, and the mathematical model of the inertial navigation algorithm is generally used.

[0140] The Kalman filtering algorithm in the application belongs to the general calculation mode in the field of inertial navigation, which will not be repeated here.

[0141] The working principle of the application is:

[0142] As shown in the formula Figure 3 The inertial navigation device and the satellite receiver are installed on the aircraft at the same time, and are fixedly connected between them. The arm between the satellite receiver and the inertial navigation device is measured in advance. The initial position information is loaded 30s before the aircraft starts to prepare to start the initial alignment, and then the aircraft takes off according to the established procedure. The initial alignment is completed during the flight process, or the inertial navigation is started and the initial alignment is started during the flight process, and the aircraft flies according to the required maneuver of the application. After the maneuver is completed, the initial alignment process is completed.

[0143] In this embodiment, the effectiveness of the application is judged by the heading angle error of the airborne strapdown inertial navigation through simulation verification, as shown inFigure 4 As shown in the figure, in the 60s rough alignment stage, the strapdown inertial navigation heading angle error reaches 6', and after the fine alignment stage takes the aircraft maneuvering method and the Kalman filter algorithm, the strapdown inertial navigation heading angle converges to less than 1' within 100s, that is, the alignment method of the application can quickly and accurately complete the alignment process.

[0144] It should be emphasized that the embodiments described in the present application are illustrative rather than restrictive, and therefore the present application includes but is not limited to the embodiments described in the specific embodiments, and any other embodiments derived by those skilled in the art according to the technical solutions of the present application also belong to the scope of protection of the present application.

Claims

1. A high-precision alignment method for airborne inertial navigation, characterized in that: The method comprises the following steps: Step 1, selecting a coarse alignment mode according to an initial alignment scene selected by a user; Step 2, performing inertial navigation coarse alignment according to the coarse alignment mode selected in step 1; Step 3, after the inertial navigation coarse alignment in step 2 is completed, a rough initial attitude angle and a position cosine matrix are obtained, and then fine alignment is performed; meanwhile, the aircraft is flown according to a set flight maneuver, if the ground alignment is selected in step 1, the aircraft only needs to start normal gliding take-off, and no special maneuver is needed; Step 4, after the aircraft maneuver is completed, the inertial navigation is completed by using a Kalman filter algorithm, the attitude angle error estimated by the Kalman filter is used to correct the navigation solution attitude angle of the strapdown inertial navigation, after the fine alignment is completed, the inertial navigation enters a navigation state, and the corrected attitude and speed information are output, and then the initial alignment process is completed; The specific method of step 2 is: (1) if the ground alignment is selected, the inertial navigation dynamic coarse alignment is performed by using a solid coordinate system coarse alignment algorithm: First, the position cosine matrix is decomposed as follows: a) solving b) solving for Recall The corresponding pose quaternion is denoted as Convert to a pose quaternion Form: When k = 0, i.e. at the start of coarse alignment coincide with the b-system, So k = 0, 1, 2,..., the quaternion update algorithm is used to calculate t k+1 the pose quaternion at time instant In the formula: Let: In the formula: Δθ = [Δθ x Δθ y Δθ z ] T — represents the gyro output angular increment in the sampling period between the time instants ((t k , t k+1 ]) in the unit sampling time, unit: rad; Normalization processing updated pose quaternion by using with the conversion relationship of t k+1 instantaneous Finally, c) solving coarse alignment start time, i.e. k = 0, When k=0, 1, 2,... In the formula: Take t n = t k+1 and t m = t k+1-N at the moment (N = 10) two points are obtained Substitute the above and into to obtain the azimuth cosine matrix; (2) if the air alignment is selected, the Beidou track angle information is input, and the coarse alignment of the Beidou track angle is performed: the angle defined as H after the synchronization of the Beidou track angle is transmitted to the inertial navigation, the initial attitude angle is zero, and then the initial attitude angle is obtained as follows: att0=[0 0H] Then, the position cosine matrix of the coarse alignment is obtained by using the attitude angle conversion position cosine matrix formula: The specific method of step 4 is: The state quantity of the Kalman filter algorithm in the fine alignment is as follows: The selected state variables include attitude angles East and north velocity errors δv of the two navigation channels x , δv y Position coordinate point errors δx, δy, δz of the two navigation channels, gyro constant drift ε x , ε y , ε z Accelerometer measurement constant bias ▽ x , ▽ y , ▽ z ; The general formula of the system observation equation is where H is a system observation matrix, is a system observation noise vector; The position of the Beidou navigation is taken as the external observation information, and the measurement equation of the observation quantity is as follows: In the formula: H v — an observation equation transfer matrix; V is an observation noise vector; where subscript Ins represents the latitude, longitude and altitude information calculated and output by the inertial navigation system, and subscript r represents the latitude, longitude and altitude information output by the satellite navigation system; H v = [0 3×6 I 3×3 0 3×6 ], and V is a 3-dimensional velocity observation white noise; At this time, the state equation of the Kalman filter is as follows: In the formula, F and G are state transition matrices and noise matrices, and the mathematical model of the inertial navigation algorithm is generally used.

2. The high-precision alignment method for airborne inertial navigation system according to claim 1, characterized in that: The specific method of step 1 is: If the ground alignment is selected, the inertial navigation coarse alignment is performed by using the solid coordinate system coarse alignment algorithm; if the air alignment is selected, the Beidou track angle information is input, and the coarse alignment of the Beidou track angle is performed.

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