A self-alignment method and device for rotary inertial navigation

Abstract

This invention relates to a self-alignment method and apparatus for a rotating inertial navigation system (INS), comprising: obtaining the orientation of a high-precision gyroscope at an initial strapdown position using a rapid coarse alignment method, and rotating the high-precision gyroscope to a predetermined east-west angle based on the orientation of the high-precision gyroscope; obtaining an initial attitude of the initial position using the orientation of the rotated high-precision gyroscope as the starting position, and rotating the high-precision gyroscope to a second position symmetrical to the initial position with the celestial axis as the central axis; performing fine alignment using filtering estimation during the rotation from the initial position to the second position and obtaining the attitude at the end of the rotation at the second position; and performing navigation tracking based on the attitude at the end of the second position and the INS data during the rotation back to the initial strapdown position around the celestial axis to complete the self-alignment. This method can fully utilize the product's high and low gyroscope combination feature, effectively ensuring the self-alignment accuracy of the rotating INS.

Description

Technical Field

[0001] This invention relates to the field of inertial navigation technology, and more specifically to a self-alignment method and apparatus for rotary inertial navigation systems. Background Technology

[0002] Currently, to improve the navigation accuracy of inertial navigation systems, efforts are being made in two ways: firstly, by designing and manufacturing higher-precision inertial components, such as using new materials, new processes, or new inertial devices; and secondly, by employing system technologies, such as advanced control strategies and algorithms. However, with the development of optical gyroscopes reaching a high level, the potential for further improvement in the accuracy of inertial devices has become limited.

[0003] In system technology, rotating inertial navigation systems (INS) typically possess self-calibration, self-alignment, and self-testing functions, making them convenient to use and maintain, and thus widely used. To reduce product size and weight, one type of rotating INS employs a combination of high and low precision gyroscopes during installation, fully utilizing the space advantage of nested gyroscope rings to achieve a high-precision, compact product.

[0004] The speed and accuracy of self-alignment are important indicators for rotating inertial navigation systems (INS). However, the widely used self-alignment methods for rotating INS currently focus on theoretical research, laboratory turntable simulation verification, or self-alignment using rotation modulation processes. There is a lack of research on high-precision self-alignment methods that combine high- and low-precision gyroscopes with the characteristics of the product itself. In particular, there is a lack of corresponding research on engineering application problems such as the use of high-precision gyroscopes and lever arm error estimation. Summary of the Invention

[0005] This invention provides a self-alignment method and apparatus for rotary inertial navigation systems, which can fully utilize the product features of high and low gyroscopes to effectively ensure the self-alignment accuracy of rotary inertial navigation systems.

[0006] On one hand, embodiments of the present invention provide a self-alignment method for a rotating inertial navigation system, characterized in that the method includes the following steps:

[0007] The orientation of the high-precision gyroscope at the initial strapdown position is obtained using a fast coarse alignment method, and the high-precision gyroscope is rotated to a set angle in the east-west direction based on the orientation of the high-precision gyroscope.

[0008] The initial attitude of the high-precision gyroscope is obtained by taking the orientation of the rotated high-precision gyroscope as the first position, and the high-precision gyroscope is rotated to a second position symmetrical to the first position with the celestial axis as the central axis.

[0009] During the rotation from the first position to the second position, a filter estimation is used for fine alignment and the attitude at the end of the rotation at the second position is obtained.

[0010] Navigation tracking is performed to complete self-alignment based on the attitude at the end of the second position and the inertial navigation data during the process of returning to the initial strapdown position around the celestial axis.

[0011] In some embodiments, the rapid coarse alignment includes second-level inertial frame alignment and parameter identification based on least-squares fitting.

[0012] In some embodiments, the rapid coarse alignment includes the steps of:

[0013] Obtain the Earth's inertial coordinate system E′ at the initial moment. I0 Attitude transformation matrix to N-system The X, Y, and Z axes of the N-system point to the east, north, and sky, respectively.

[0014] according to and f B Obtain the current B-frame relative to the initial strapdown inertial navigation system B-frame. I0 rotation matrix and the B I0 Specific velocity in the system Wherein, the B system is the body coordinate system of the inertial measurement unit, with its X-axis pointing upwards and its Y and Z axes horizontal. f is the projection of the angular velocity of frame B relative to the geocentric inertial coordinate system I into frame B. B The accelerometer of the inertial measurement unit (IMU) outputs the specific force in the body coordinate system B, where the origin of the I system is located at the center of the Earth and the coordinate axis directions remain unchanged in inertial space.

[0015] Based on the above and stated Obtain the B I0 Reset to Earth's inertial coordinate system E′ at the initial moment I0 pose transformation matrix

[0016] According to the above The and stated Calculate the initial attitude transformation matrix of the inertial measurement unit (IMU).

[0017] In some embodiments, the attitude transformation matrix Based on the first formula, which includes:

[0018]

[0019] Where, ω IE L0 is the Earth's rotational angular velocity, L0 is the latitude of the rotating inertial navigation system at the initial alignment time t0, and t is the current alignment time.

[0020] In some embodiments, the and the This is obtained based on the second formula, which includes:

[0021]

[0022]

[0023]

[0024] Where I is the identity matrix, τ is a variable in the integral expression, t is the current alignment time, and t0 is the alignment start time.

[0025] In some embodiments, the Obtained based on a third formula, which includes:

[0026]

[0027]

[0028] t1 = 1 / 2 * t2,

[0029] Among them, f N For the N-series, for transpose, for When t0 is 0 and t is the value of t1. for When t0 is 0 and t is the value of t2. for When t0 is 0 and t is the value of t1. for τ is a variable in the integral expression where t0 takes the value 0 and t takes the value t2.

[0030] In some embodiments, the Based on the fourth formula, which includes:

[0031]

[0032]

[0033]

[0034] Among them, among them, Let the direction cosine matrix be from system B to system N. Let L be the direction cosine matrix from the initial coordinate system B0 to the navigation system L at time zero. The direction cosine matrix is ​​from B system to N system, and the X, Y and Z axes of the navigation system L system point to the north, east and ground respectively.

[0035] In some embodiments, obtaining the orientation of the high-precision gyroscope at the initial strapdown position using a rapid coarse alignment method includes the following steps:

[0036] Using the initial attitude matrix and storage preset time and f B The attitude matrix at each time step is obtained by performing navigation calculations. and velocity v N (t);

[0037] Using a quadratic polynomial to describe the velocity v N (t) The coefficients X of the eastward and northward velocity fitting are obtained by fitting the data. E and X N And based on the coefficient X E and X N Calculate the attitude error corresponding to the preset time;

[0038] The attitude error is corrected, and the orientation of the high-precision gyroscope is obtained based on the corrected attitude matrix.

[0039] In some embodiments, the step of using filtered estimation for fine alignment includes the following steps:

[0040] The state equation is established based on the first formula, which includes:

[0041]

[0042]

[0043]

[0044]

[0045]

[0046]

[0047]

[0048]

[0049] Where F is the state transition matrix, which is a 20×20 dimensional matrix, and w is the system state noise, which is a 20×1 dimensional vector; x INSSys Let x be the state vector of the inertial navigation system parameter error. INSSens Let x be the accelerometer and gyroscope parameter error state vector.LB Let be the state vector of the inner arm. These represent the components of the position error angle of the downloaded volume in the navigation coordinate system in the X, Y, and Z axes, respectively. These represent the components of the navigation system's velocity error in the X, Y, and Z axes, respectively. These represent the components of the navigation system's attitude error angle in the X, Y, and Z axes, respectively; δh represents the carrier's height error; and x... INSSens It contains three accelerometer zero bias errors B' and three corresponding gyroscope zero bias errors B″, x LB This includes, in sequence, the projections of the accelerometers on the non-rotational axes onto each horizontal axis. Let X be the components of the displacement angular velocity in the X, Y, and Z directions under the navigation system. R represents the components of the Earth's rotational angular velocity in the Y and Z directions in the navigation system. M R is the radius of curvature of the meridian. N Let L be the radius of curvature of the zonal parallel, L be the latitude, and h be the altitude. Let X represent the components of velocity in the X, Y, and Z directions within the navigation frame. These are the specific forces measured by the three accelerometers in the navigation frame, respectively. skew(x) represents the antisymmetric matrix of vector x, and the subscripts are... (l-1) Δt represents the value of the previous sampling period, and Δt represents the sampling interval.

[0050] The measurement equation is established based on the second formula, which includes:

[0051] z = Hx + v

[0052] H = [I 3×3 0 3×17 ],

[0053]

[0054] Where Lat is the latitude of the inertial navigation system (INS) solution, λ is the longitude of the INS solution, and h is the altitude of the INS solution. iniCon λ is the latitude of the local location. iniCon h is the longitude of the local location. iniCon Where is the altitude of the local point, and v is the measurement noise;

[0055] Kalman filtering is performed based on the state equation and the measurement equation to correct inertial navigation system parameter errors and device parameter errors in real time.

[0056] Secondly, embodiments of the present invention also provide a self-alignment device for a rotating inertial navigation system, characterized in that it comprises,

[0057] The coarse alignment module is used to obtain the orientation of the high-precision gyroscope in the initial strapdown position using a fast coarse alignment method, and to rotate the high-precision gyroscope to a set angle in the east-west direction according to the orientation of the high-precision gyroscope.

[0058] A rotation module is used to obtain the initial attitude of the first position with the orientation of the high-precision gyroscope after rotation as the first position, and rotate the high-precision gyroscope to a second position symmetrical to the first position with the celestial axis as the central axis.

[0059] A fine alignment module is used to perform fine alignment by using filtering estimation during the rotation from the first position to the second position and to obtain the attitude at the end of the rotation at the second position.

[0060] The navigation tracking module is used to perform navigation tracking and complete self-alignment based on the attitude at the end of the second position and the inertial navigation data during the process of rotating around the celestial axis to the initial strapdown position.

[0061] This invention provides a self-alignment method for a rotating inertial navigation system (INS) using a combination of high- and low-precision gyroscopes. The method employs rapid coarse alignment (optionally on the order of seconds) to achieve coarse north finding, then rotates the high-precision gyroscope to near the east-west direction, fully utilizing the high-precision gyroscope for self-alignment. Simultaneously, alignment data is reused at the initial east-west position, followed by fine alignment using two symmetrical positions. By fully utilizing the product's high- and low-precision gyroscope combination, the self-alignment accuracy of the rotating INS is effectively guaranteed, and data reuse improves the effective alignment time. Compared with related technologies, the advantage of this invention is that it provides a self-alignment method that does not rely on external information, fully utilizing the product's high- and low-precision gyroscope combination to effectively improve self-alignment accuracy. Attached Figure Description

[0062] To more clearly illustrate the technical solutions in the embodiments of the present invention, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0063] Figure 1 This is a schematic flowchart of a self-alignment method for a rotating inertial navigation system provided in an embodiment of the present invention.

[0064] Figure 2 This is a schematic flowchart of a self-alignment method for a rotating inertial navigation system provided in an embodiment of the present invention.

[0065] Figure 3 This is a top view schematic diagram of the self-alignment process provided in an embodiment of the present invention;

[0066] Figure 4 This is a schematic diagram of a self-alignment device for a rotating inertial navigation system provided in an embodiment of the present invention. Detailed Implementation

[0067] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0068] like Figure 1 As shown, this embodiment of the invention provides a self-alignment method for a rotating inertial navigation system, the method comprising the following steps:

[0069] S100: The orientation of the high-precision gyroscope at the initial strapdown position is obtained by using a fast coarse alignment method, and the high-precision gyroscope is rotated to a set angle in the east-west direction according to the orientation of the high-precision gyroscope.

[0070] S200: Obtain the initial attitude of the first position by taking the orientation of the high-precision gyroscope after rotation as the first position, and rotate the high-precision gyroscope to a second position symmetrical to the first position with the celestial axis as the central axis;

[0071] S300: During the rotation from the first position to the second position, a filter estimation is used for fine alignment and the attitude at the end of the rotation at the second position is obtained.

[0072] S400: Navigation tracking is performed to complete self-alignment based on the attitude at the end of the second position and the inertial navigation data during the process of returning to the initial strapdown position around the celestial axis.

[0073] It should be noted that, typically on a carrier that needs to be aligned, the inertial measurement unit consists of three gyroscopes and three accelerometers. The initial strapdown position refers to the position where the rotating inertial navigation system is switched to zero.

[0074] This invention provides a self-alignment method for a rotating inertial navigation system (INS) using a combination of high- and low-precision gyroscopes. The method employs rapid coarse alignment (optionally on the order of seconds) to achieve coarse north finding, then rotates the high-precision gyroscope to near the east-west direction, fully utilizing the high-precision gyroscope for self-alignment. Simultaneously, alignment data is reused at the initial east-west position, followed by fine alignment using two symmetrical positions. By fully utilizing the product's high- and low-precision gyroscope combination, the self-alignment accuracy of the rotating INS is effectively guaranteed, and data reuse improves the effective alignment time. Compared with related technologies, the advantage of this invention is that it provides a self-alignment method that does not rely on external information, fully utilizing the product's high- and low-precision gyroscope combination to effectively improve self-alignment accuracy.

[0075] Preferably, rapid coarse alignment includes second-level inertial frame alignment and parameter identification based on least-squares fitting.

[0076] Furthermore, the rapid coarse alignment in S100 includes the following steps:

[0077] S101: Obtain the initial Earth inertial coordinate system E′ I0 Attitude transformation matrix to N-system The X, Y, and Z axes of the N-system point to the east, north, and sky, respectively.

[0078] S102: According to and f B Obtain the current B-frame relative to the initial strapdown inertial navigation system B-frame. I0 rotation matrix and the B I0 Specific velocity in the system Wherein, the B system is the body coordinate system of the inertial measurement unit, with its X-axis pointing upwards and its Y and Z axes horizontal. f is the projection of the angular velocity of frame B relative to the geocentric inertial coordinate system I into frame B. B The accelerometer of the inertial measurement unit (IMU) outputs the specific force in the body coordinate system B, where the origin of the I system is located at the center of the Earth and the coordinate axis directions remain unchanged in inertial space.

[0079] S103: Based on the above and stated Obtain the B I0 Reset to Earth's inertial coordinate system E′ at the initial moment I0 pose transformation matrix

[0080] S104: According to the above The and stated Calculate the initial attitude transformation matrix of the inertial measurement unit (IMU).

[0081] It should be noted that the force ratio is the inertial reference acceleration minus the gravitational acceleration.

[0082] In some embodiments, the pose transformation matrix Based on the first formula, which includes:

[0083]

[0084] Where, ω IE L0 is the Earth's rotational angular velocity, L0 is the latitude of the rotating inertial navigation system at the initial alignment time t0, and t is the current alignment time.

[0085] In some embodiments, and the This is obtained based on the second formula, which includes:

[0086]

[0087]

[0088]

[0089] Where I is the identity matrix, τ is a variable in the integral expression, t is the current alignment time, and t0 is the alignment start time.

[0090] It should be noted that in the second formula Indicates to The update, the left side of the equals sign is the updated version. The right side of the equals sign is the original value before the update. express The cross product.

[0091] In some embodiments, This is obtained based on the third formula, which includes:

[0092]

[0093]

[0094] t1 = 1 / 2 * t2,

[0095] Among them, f N For the N-series, for transpose, for When t0 is 0 and t is the value of t1. for When t0 is 0 and t is the value of t2. for When t0 is 0 and t is the value of t1. for τ is a variable in the integral expression where t0 takes the value 0 and t takes the value t2.

[0096] In some embodiments, This is obtained based on the fourth formula, which includes:

[0097]

[0098]

[0099]

[0100] in, Let the direction cosine matrix be from system B to system N. Let L be the direction cosine matrix from the initial coordinate system B0 to the navigation system L at time zero. The direction cosine matrix is ​​from the B system to the N system, and the X, Y, and Z axes in the navigation system L point to the north, east, and ground, respectively.

[0101] It should be noted that the B0 system is the coordinate system at the initial zero time, so It is equal to When t takes the value 0, that is

[0102] It is understandable that the parameters to be sought are the initial attitude matrix (direction cosine matrix). Since the L-frame points to North, East, and Earth, and the N-frame points to East, North, and Sky, and the two have a fixed relationship, they can be calculated as needed. or

[0103] In some embodiments, S100 includes the step of:

[0104] S105: Using the initial attitude matrix and storage preset time and f B The attitude matrix at each time step is obtained by performing navigation calculations. and velocity v N (t);

[0105] S106: Using a quadratic polynomial to measure the velocity v N (t) The coefficients X of the eastward and northward velocity fitting are obtained by fitting the data. E and X N And based on the coefficient X E and X N Calculate the attitude error corresponding to the preset time;

[0106] S107: Correct the attitude error and obtain the orientation of the high-precision gyroscope based on the corrected attitude matrix.

[0107] It should be noted that the preset time can be selected from a value between 10 and 20 seconds, and can be preferably set to 10 seconds. If the preset time is 10 seconds, then the attitude error calculated in step S106 is the attitude error at the 10th second.

[0108] In some embodiments, the fine alignment using filtered estimation in S300 includes the following steps:

[0109] S301: Establish the state equation based on the first formula, wherein the first formula includes:

[0110]

[0111]

[0112]

[0113]

[0114]

[0115]

[0116]

[0117]

[0118] Where F is the state transition matrix, which is a 20×20 dimensional matrix, and w is the system state noise, which is a 20×1 dimensional vector; x INSSys Let x be the state vector of the inertial navigation system parameter error. INSSens Let x be the accelerometer and gyroscope parameter error state vector. LB Let be the state vector of the inner arm. These represent the components of the position error angle of the downloaded volume in the navigation coordinate system in the X, Y, and Z axes, respectively. These represent the components of the navigation system's velocity error in the X, Y, and Z axes, respectively. These represent the components of the navigation system's attitude error angle in the X, Y, and Z axes, respectively; δh represents the carrier's height error; and x... INSSens It includes, in sequence, the zero bias errors of three accelerometers B' and three gyroscopes B″ on the X, Y, and Z axes, respectively. LB This includes, in sequence, the projections of the accelerometers on the non-rotational axes onto each horizontal axis. Let X be the components of the displacement angular velocity in the X, Y, and Z directions under the navigation system. R represents the components of the Earth's rotational angular velocity in the Y and Z directions in the navigation system. M R is the radius of curvature of the meridian. N Let L be the radius of curvature of the zonal parallel, L be the latitude, and h be the altitude. Let X represent the components of velocity in the X, Y, and Z directions within the navigation frame. These are the specific forces measured by the three accelerometers in the navigation frame, respectively. skew(x) represents the antisymmetric matrix of vector x, and the subscripts are... (l-1) Δt represents the value of the previous sampling period, and Δt represents the sampling interval.

[0119] S302: Establish the measurement equation based on the second formula, which includes:

[0120] z = Hx + v

[0121] H = [I 3×3 0 3×17 ],

[0122]

[0123] Where Lat is the latitude of the inertial navigation system (INS) solution, λ is the longitude of the INS solution, and h is the altitude of the INS solution. iniCon λ is the latitude of the local location. iniCon h is the longitude of the local location. iniCon Where is the altitude of the local point, and v is the measurement noise;

[0124] S303: Perform Kalman filtering based on the state equation and the measurement equation to correct inertial navigation system parameter errors and device parameter errors in real time.

[0125] This invention employs a second-level rapid coarse alignment to achieve coarse north finding, followed by rotating a high-precision gyroscope to near the east-west direction. This fully utilizes the high-precision gyroscope for self-alignment. Simultaneously, alignment data is reused at the initial east-west position, and then fine alignment is performed using two symmetrical positions. The process employs a full-dimensional navigation parameter model and a lever arm estimation model, utilizing a position combination method. By fully leveraging the product's high- and low-precision gyroscope combination, the self-alignment accuracy of the rotating inertial navigation system is effectively guaranteed, while data reuse improves the effective alignment time. Compared with related technologies, the advantages of this invention lie in providing a self-alignment method that does not rely on external information, fully utilizing the product's high- and low-precision gyroscope combination, and employing a data reuse strategy and lever arm error modeling and estimation method to effectively improve self-alignment accuracy.

[0126] like Figure 2 As shown, in a specific embodiment, the projection of the angular velocity of the inertial measurement unit (IMU) gyroscope output body coordinate system B relative to the geocentric inertial coordinate system I in the B system is ω. I BB The IMU's accelerometer outputs the specific force f in the body coordinate system B. B In the geocentric inertial coordinate system I, the origin is located at the center of the Earth, and the coordinate axes remain unchanged in inertial space. In the navigation system N, the X, Y, and Z axes point east, north, and sky, respectively. In the navigation system L, the X, Y, and Z axes point north, east, and ground, respectively. In the body coordinate system B, the X axis points skyward, and the Y and Z axes are horizontal. A self-alignment method for rotating inertial navigation can be developed through steps S1 to S6:

[0127] S1: Perform rapid coarse alignment within seconds at the initial strapdown position and determine the orientation of the high-precision gyroscope at the initial strapdown position. S1 may specifically include steps S11 to S17:

[0128] S11: Obtain the initial Earth inertial coordinate system E′ I0 Attitude transformation matrix to N-system According to the formula:

[0129] Sure

[0130] Where L0 is the latitude of the rotating inertial navigation system aligned to the initial time t0, and ω IE t represents the Earth's rotational angular velocity, and t represents the current time.

[0131] S12: According to and f B Obtain the current B-frame relative to the initial strapdown inertial navigation system B-frame. I0 rotation matrix and the initial strapdown inertial navigation inertial coordinate system B I0 The specific velocity below

[0132] Step S12 specifically includes S121 to S122:

[0133] S121: According to Attitude update via strapdown inertial navigation Calculate the current B frame relative to the strapdown inertial navigation system B at the initial time. I0 rotation matrix in, I is the identity matrix;

[0134] S122: According to the formula The initial strapdown inertial navigation system coordinate system B was calculated. I0 The specific velocity below τ is a variable in the integral expression.

[0135] It should be noted that in this embodiment, specific velocity is the integral of specific force. Specific force is a unit of acceleration, and its integral is velocity, which is specific velocity.

[0136] S13: Based on and Obtain the initial strapdown inertial navigation system coordinate system B. I0 Earth inertial coordinate system E′ at the initial moment I0 pose transformation matrix

[0137] Step S13 may specifically include S131 to S132:

[0138] S131: According to the formula The initial Earth inertial coordinate system E′ was calculated. I0 The specific velocity below Among them, f N For the N-series, for The transpose, i.e.

[0139] S132: Based on the strapdown inertial navigation inertial coordinate system B at the initial moment I0 The specific velocity below and the Earth's inertial coordinate system E′ at the initial moment I0 The specific velocity below According to the formula

[0140]

[0141] The initial time of the strapdown inertial navigation system's inertial coordinate system BI0 and the initial time of the Earth's inertial coordinate system E′ were calculated. I0 pose transformation matrix

[0142] in, for The integral of t0 is 0 and t is t1. The integral is defined as t0 taking 0 and t taking t2; for The integral of t0 is 0 and t is t1. Let t0 be 0 and t be t2, where t1 = 1 / 2 * t2, and τ is a variable in the integral expression.

[0143] In this embodiment, t2 is taken as the current time t.

[0144] S14: According to and Obtain the initial attitude transformation matrix of the IMU. Step S14 may specifically include S141 to S142:

[0145] S141: According to the formula Determine the Earth's inertial coordinate system at the initial moment. E Attitude transformation matrix from I0 to B system

[0146] S142: According to the formula The initial attitude transformation matrix of the IMU at the initial time (time 0) at this position is calculated. and

[0147] S15: After the alignment time is completed as described above, use the initial attitude matrix. And store for 10 seconds and f B The attitude matrix at each time step is obtained by performing navigation calculations. and velocity v N (t), the coefficients X of the eastward and northward velocity fitting are obtained by using quadratic polynomial fitting. E and X N (The fitting coefficients include constant, first-order, and second-order coefficients).

[0148] S16: Calculate the attitude error at time t (10s):

[0149] Among them, X E (1) represents the coefficient of the first-order term in the quadratic polynomial fitting, X N (2) represents the coefficients of the quadratic term in the fitting of the quadratic polynomial.

[0150] S17: Correct the attitude error at time t:

[0151] φ = norm(φ L )

[0152]

[0153] Here, norm(x) represents the magnitude of vector x.

[0154] The orientation of the high-precision gyroscope (Y-gyroscope) is obtained based on the corrected attitude matrix.

[0155] S2: Using the high-precision gyroscope orientation calculated in step S1, rotate the high-precision gyroscope to a set angle near the east-west direction.

[0156] It should be noted that in this embodiment of the invention, the angle set near the east-west direction can be set to 3 to 10 degrees, typically 5 degrees east of north. For example... Figure 3 As shown, the angle θ between the orientation of the high-precision gyroscope (Y gyroscope) obtained in step S1 and the set east-west angle (e.g., 5 degrees east of north) is calculated. Figure 3The angle θ in the equation is then used to rotate the inertial measurement unit around the vertical axis (-X axis) by an angle θ to a set east-west angle (e.g., θ angle). Figure 3 (The rotation corresponding to sequence number 1 used to identify the process sequence).

[0157] S3: Starting from the east-west orientation after the rotation in step S2, begin the coarse alignment process for fine alignment between the two positions, pause for t seconds, and obtain the attitude at the initial moment of the first position. Simultaneously, cache the initial position inertial navigation data. For S3, refer to S1 for specific inertial frame alignment, according to the formula... The initial attitude transformation matrix of the IMU at the initial time (time 0) at this position is calculated. Simultaneously cache the inertial navigation data that remains for t seconds during step S3 (including... and f B ).

[0158] S4: After staying in the first position for t seconds, rotate 180 degrees around the celestial axis at a speed of w to the second position, and stay there for t seconds as well.

[0159] It should be noted that in this embodiment of the invention, t represents the dwell time at the two positions, which can be set according to the total self-alignment time. For example, if the total self-alignment time requirement is 300 seconds, then after deducting the time for east-west rotation and 180-degree rotation, the dwell time at each position is approximately 120 seconds, so t can be set to 120 seconds. w represents the rotation speed, which can be set according to the capability of the indexing mechanism, generally set to 20 degrees / second or 30 degrees / second. The process of rotating 180 degrees around the axial direction at a speed of w to reach the second position is as follows: Figure 3 As shown in number 2 in the diagram.

[0160] S5: Simultaneously with the start of step S4, using the established Kalman filter model, perform full-range filtering estimation starting from the first position and continuing until the end at the second position, obtaining the attitude at the end of the second position. Step S5 may specifically include S51 to S52:

[0161] S51: Establish the state equations.

[0162] The state equation of the system is:

[0163]

[0164] Where F is the state transition matrix, a 20×20 dimensional matrix, and w is the system state noise, a 20×1 dimensional vector; the state vector The x INSSys Let x be the parameter error state vector of the inertial navigation system. INSSens The x is the error state vector of the accelerometer and gyroscope parameters. LBLet x be the state vector of the inner arm. In one embodiment of the invention, the state vector x is a 20-dimensional vector, containing a 10-dimensional inertial navigation system parameter error state vector x. INSSys 6-dimensional device parameter error state vector x INSSens And the 4D lever arm state vector.

[0165] Where x INSSys Specifically, it is expressed as follows:

[0166]

[0167] These represent the components of the position error angle of the downloaded volume in the navigation coordinate system in the X, Y, and Z axes, respectively. These represent the components of the navigation system's velocity error in the X, Y, and Z axes, respectively. δ0 represents the components of the attitude error angle of the navigation system in the X, Y, and Z axes, respectively, and δh represents the height error of the carrier.

[0168] x INSSens It contains three accelerometer zero bias errors B' and three gyroscope zero bias errors B″.

[0169] x LB It sequentially includes the projections of the accelerometers on the non-rotating axes onto each horizontal axis.

[0170] F is the state transition matrix, where F 11 The coefficient matrix of the linearized state error equation for a 10×10 dimensional strapdown inertial navigation system is expressed as follows:

[0171]

[0172] in, Let X be the components of the displacement angular velocity in the X, Y, and Z directions under the navigation system. R represents the components of the Earth's rotational angular velocity in the Y and Z directions in the navigation system. M R is the radius of curvature of the meridian. N Let L be the radius of curvature of the zonal parallel, L be the latitude, and h be the altitude. Let X represent the components of velocity in the X, Y, and Z directions within the navigation frame. These are the specific force values ​​measured by three accelerometers in the navigation system;

[0173] in This is the attitude matrix calculated by the inertial navigation system.

[0174] in

[0175] skew(x) denotes the antisymmetric matrix of vector x, with subscripts... (l-1) Δt represents the value of the previous sampling period, and Δt represents the sampling interval.

[0176] S52: Establish measurement equations.

[0177] z = Hx + v, where the measurement matrix H = [I 3×3 0 3×17 ],

[0178] Observation

[0179] Where Lat is the latitude of the inertial navigation system (INS) solution, λ is the longitude of the INS solution, and h is the altitude of the INS solution. iniCon λ is the latitude of the local location. iniCon h is the longitude of the local location. iniCon v represents the altitude of the local location, and v represents the measurement noise.

[0180] S53: Using the state equation and measurement equation, perform Kalman filtering to correct inertial navigation system parameter errors and device parameter errors in real time, and complete the precise alignment.

[0181] After obtaining the state equation and measurement equation of the navigation system, a classic Kalman filter feedback correction method is adopted, specifically as follows:

[0182] Extrapolation calculation period:

[0183]

[0184]

[0185] Update calculation cycle:

[0186]

[0187]

[0188]

[0189] Feedback correction:

[0190]

[0191] It should be noted that (+E) is a superscript, indicating that after the state update in the basic equation of the discrete Kalman filter, the specified parameter at the specified time (here, t) is updated. nThe value of the parameter at a given time is (+C) (a superscript indicating the value of the parameter at a given time after the control equation in the basic equation of the discrete Kalman filter); the value of the parameter at a given time is (-) (a superscript indicating the value of the parameter at a given time before the state update equation and the control equation take effect).

[0192] Understandably, after Kalman filtering, estimates of each error state can be obtained. These estimates are then used to correct the navigation parameters, which are then used as the initial values ​​for the next navigation iteration to continue the navigation calculation until the process ends.

[0193] S6: After the second position dwell time ends, rotate around the celestial axis back to the initial strapdown position, and simultaneously use inertial navigation data to navigate back to the initial position, outputting the high-precision attitude of the strapdown position, and perform navigation tracking to complete self-alignment.

[0194] In this embodiment of the invention, the final attitude given is the initial strapdown position. Therefore, after completing the fine alignment at the second position, it is necessary to rotate back to the initial strapdown position around the celestial axis (e.g., ...). Figure 3 (The process corresponding to number 3 in the middle) can be used to perform navigation calculations by utilizing the output of the inertial navigation system.

[0195] like Figure 4 As shown, this embodiment of the invention also provides a self-alignment device for a rotating inertial navigation system, characterized in that it includes,

[0196] The coarse alignment module is used to obtain the orientation of the high-precision gyroscope in the initial strapdown position using a fast coarse alignment method, and to rotate the high-precision gyroscope to a set angle in the east-west direction according to the orientation of the high-precision gyroscope.

[0197] A rotation module is used to obtain the initial attitude of the first position with the orientation of the high-precision gyroscope after rotation as the first position, and rotate the high-precision gyroscope to a second position symmetrical to the first position with the celestial axis as the central axis.

[0198] A fine alignment module is used to perform fine alignment by using filtering estimation during the rotation from the first position to the second position and to obtain the attitude at the end of the rotation at the second position.

[0199] The navigation tracking module is used to perform navigation tracking and complete self-alignment based on the attitude at the end of the second position and the inertial navigation data during the process of rotating around the celestial axis to the initial strapdown position.

[0200] In some embodiments, rapid coarse alignment includes second-level inertial frame alignment and parameter identification based on least-squares fitting.

[0201] In some embodiments, the coarse alignment module is also used for:

[0202] Obtain the Earth's inertial coordinate system E′ at the initial moment.I0 Attitude transformation matrix to N-system The X, Y, and Z axes of the N-system point to the east, north, and sky, respectively.

[0203] according to and f B Obtain the current B-frame relative to the initial strapdown inertial navigation system B-frame. I0 rotation matrix and the B I0 Specific velocity in the system Wherein, the B system is the body coordinate system, with its X-axis pointing upwards, and its Y and Z axes being horizontal. The projection of the angular velocity of the inertial measurement unit (IMU) gyroscope output body coordinate system B relative to the geocentric inertial coordinate system I in the B system, wherein f B The accelerometer of the inertial measurement unit (IMU) outputs the specific force in the body coordinate system B, where the origin of the I system is located at the center of the Earth and the coordinate axis directions remain unchanged in inertial space.

[0204] Based on the above and stated Obtain the B I0 Reset to Earth's inertial coordinate system E′ at the initial moment I0 pose transformation matrix

[0205] According to the above The and stated Calculate the initial attitude transformation matrix of the inertial measurement unit (IMU).

[0206] In some embodiments, the pose transformation matrix Based on the first formula, which includes:

[0207]

[0208] Where, ω IE L0 is the Earth's rotational angular velocity, L0 is the latitude of the rotating inertial navigation system at the initial alignment time t0, and t is the current alignment time.

[0209] In some embodiments, and the This is obtained based on the second formula, which includes:

[0210]

[0211]

[0212]

[0213] Where I is the identity matrix, τ is a variable in the integral expression, t is the current alignment time, and t0 is the alignment start time.

[0214] In some embodiments, Obtained based on a third formula, which includes:

[0215]

[0216]

[0217] t1 = 1 / 2 * t2,

[0218] Among them, f N For the N-series, for transpose, for When t0 is 0 and t is the value of t1. for When t0 is 0 and t is the value of t2. for When t0 is 0 and t is the value of t1. for τ is a variable in the integral expression where t0 takes the value 0 and t takes the value t2.

[0219] In some embodiments, This is obtained based on the fourth formula, which includes:

[0220]

[0221]

[0222]

[0223] Among them, among them, Let the direction cosine matrix be from system B to system N. Let L be the direction cosine matrix from coordinate system B0 to coordinate system L at the initial time zero. Let be the direction cosine matrix from the B system to the N system.

[0224] In some embodiments, the coarse alignment module is also used for:

[0225] Using the initial attitude matrix and storage preset time and f B The attitude matrix at each time step is obtained by performing navigation calculations. and velocity v N (t);

[0226] Using a quadratic polynomial to describe the velocity vN (t) The coefficients X of the eastward and northward velocity fitting are obtained by fitting the data. E and X N And based on the coefficient X E and X N Calculate the attitude error corresponding to the preset time;

[0227] The attitude error is corrected, and the orientation of the high-precision gyroscope is obtained based on the corrected attitude matrix.

[0228] In some embodiments, the alignment module is also used for:

[0229] The state equation is established based on the first formula, which includes:

[0230]

[0231]

[0232]

[0233]

[0234]

[0235]

[0236]

[0237]

[0238] Where F is the state transition matrix, which is a 20×20 dimensional matrix, and w is the system state noise, which is a 20×1 dimensional vector; x INSSys Let x be the state vector of the inertial navigation system parameter error. INSSens Let x be the accelerometer and gyroscope parameter error state vector. LB Let be the state vector of the inner arm. These represent the components of the position error angle of the downloaded volume in the navigation coordinate system in the X, Y, and Z axes, respectively. These represent the components of the navigation system's velocity error in the X, Y, and Z axes, respectively. These represent the components of the navigation system's attitude error angle in the X, Y, and Z axes, respectively; δh represents the carrier's height error; and x... INSSens It contains three accelerometer zero-bias errors B' and three gyroscope zero-bias errors B″, x LB This includes, in sequence, the projections of the accelerometers on the non-rotational axes onto each horizontal axis. Let X be the components of the displacement angular velocity in the X, Y, and Z directions under the navigation system. R represents the components of the Earth's rotational angular velocity in the Y and Z directions in the navigation system. M R is the radius of curvature of the meridian. N Let L be the radius of curvature of the zonal parallel, L be the latitude, and h be the altitude. Let X represent the components of velocity in the X, Y, and Z directions within the navigation frame. These are the specific forces measured by the three accelerometers in the navigation frame, respectively. skew(x) represents the antisymmetric matrix of vector x, and the subscripts are... (l-1) Δt represents the value of the previous sampling period, and Δt represents the sampling interval.

[0239] The measurement equation is established based on the second formula, which includes:

[0240] z = Hx + v

[0241] H = [I 3×3 0 3×17 ],

[0242]

[0243] Where Lat is the latitude of the inertial navigation system (INS) solution, λ is the longitude of the INS solution, and h is the altitude of the INS solution. iniCon λ is the latitude of the local location. iniCon h is the longitude of the local location. iniCon Where is the altitude of the local point, and v is the measurement noise;

[0244] Kalman filtering is performed based on the state equation and the measurement equation to correct inertial navigation system parameter errors and device parameter errors in real time.

[0245] Those skilled in the art will understand that all or some of the steps, systems, and apparatuses disclosed above, and their functional modules / units, can be implemented as software, firmware, hardware, or suitable combinations thereof. In hardware implementations, the division between functional modules / units mentioned above does not necessarily correspond to the division of physical components; for example, a physical component may have multiple functions, or a function or step may be performed collaboratively by several physical components. Some or all physical components may be implemented as software executed by a processor, such as a central processing unit, digital signal processor, or microprocessor, or as hardware, or as an integrated circuit, such as an application-specific integrated circuit (ASIC). Such software can be distributed on a computer-readable storage medium, which may include computer-readable storage media (or non-transitory media) and communication media (or transient media).

[0246] It should be noted that in this invention, relational terms such as "first" and "second" are used merely to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0247] The above description is merely a specific embodiment of the present invention, enabling those skilled in the art to understand or implement the invention. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the invention. Therefore, the present invention is not to be limited to the embodiments shown herein, but is to be accorded the widest scope consistent with the principles and novel features claimed herein.

Claims

1. A self-alignment method for a rotating inertial navigation system, characterized in that, The method includes the following steps: The orientation of the high-precision gyroscope at the initial strapdown position is obtained using a fast coarse alignment method, and the high-precision gyroscope is rotated to a set angle in the east-west direction based on the orientation of the high-precision gyroscope. The initial attitude of the high-precision gyroscope is obtained by taking the orientation of the rotated high-precision gyroscope as the first position, and the high-precision gyroscope is rotated to a second position symmetrical to the first position with the celestial axis as the central axis. During the rotation from the first position to the second position, a filter estimation is used for fine alignment and the attitude at the end of the rotation at the second position is obtained. Navigation tracking is performed to complete self-alignment based on the attitude at the end of the second position and the inertial navigation data during the process of returning to the initial strapdown position around the celestial axis. The rapid coarse alignment includes the following steps: Obtain the Earth's inertial coordinate system at the initial moment Attitude transformation matrix to N-system The X, Y, and Z axes of the N-system point to the east, north, and sky, respectively. according to and Obtain the current B-frame relative to the initial strapdown inertial navigation system coordinate system. rotation matrix and the Specific velocity in the system Wherein, the B system is the body coordinate system of the inertial measurement unit, with its X-axis pointing upwards and its Y and Z axes horizontal. The projection of the angular velocity of frame B relative to the geocentric inertial coordinate system I into frame B is the following. The accelerometer of the inertial measurement unit (IMU) outputs the specific force in the body coordinate system B, where the origin of the I system is located at the center of the Earth and the coordinate axis directions remain unchanged in inertial space. Based on the above and stated Obtain the Reset to the Earth's inertial coordinate system at the initial moment pose transformation matrix ; According to the above The above and stated Calculate the initial attitude transformation matrix of the inertial measurement unit (IMU). .

2. The self-alignment method for a rotating inertial navigation system as described in claim 1, characterized in that, The rapid coarse alignment includes second-level inertial frame alignment and parameter identification based on least-squares fitting.

3. The self-alignment method for a rotating inertial navigation system as described in claim 2, characterized in that, The attitude transformation matrix Based on the first formula, which includes: ， in, The angular velocity of Earth's rotation. To align with the start time Latitude of rotating inertial navigation system To align with the current time.

4. The self-alignment method for a rotating inertial navigation system as described in claim 2, characterized in that, The and the This is obtained based on the second formula, which includes: ， ， ， Among them, the It is the identity matrix. For variables in the integral expression, For the current alignment time, To align with the start time.

5. A self-alignment method for a rotating inertial navigation system as described in claim 2, characterized in that, The Obtained based on a third formula, which includes: ， ， t1 = 1 / 2 * t2, in, For the N-series, for transpose, for When t0 is 0 and t is the value of t1. for When t0 is 0 and t is the value of t2. for When t0 is 0 and t is the value of t1. for When t0 is 0 and t is the value of t2. For variables in the integral expression.

6. The self-alignment method for a rotating inertial navigation system as described in claim 2, characterized in that, The Based on the fourth formula, which includes: ， ， ， Among them, among them, Let the direction cosine matrix be from system B to system N. Let L be the direction cosine matrix from the initial coordinate system B0 to the navigation system L at time zero. The direction cosine matrix is ​​from the B system to the N system. In the navigation system L, the X, Y, and Z axes point to the north, east, and ground, respectively.

7. A self-alignment method for a rotating inertial navigation system as described in any one of claims 2 to 6, characterized in that, The method of obtaining the orientation of the high-precision gyroscope at the initial strapdown position using a fast coarse alignment method includes the following steps: Using the initial attitude matrix and storage preset time and The attitude matrix at each time step is obtained by performing navigation calculations. and speed ; Using a quadratic polynomial to describe the velocity The coefficients for fitting the eastward and northward velocities are obtained by performing a fitting operation. and and based on the coefficients and Calculate the attitude error corresponding to the preset time; The attitude error is corrected, and the orientation of the high-precision gyroscope is obtained based on the corrected attitude matrix.

8. The self-alignment method for a rotating inertial navigation system as described in claim 1, characterized in that, The step of using filtered estimation for fine alignment includes the following steps: The state equation is established based on the first formula, which includes: ， ， ， in, Let be the state transition matrix and be A 3D matrix The system state noise is and is A dimensional vector; This is the state vector of parameter errors in the inertial navigation system. This is the error state vector for accelerometer and gyroscope parameters. Let be the state vector of the inner arm. , , These represent the components of the position error angle of the downloaded volume in the navigation coordinate system in the X, Y, and Z axes, respectively. , , These represent the components of the navigation system's velocity error in the X, Y, and Z axes, respectively. , , These represent the components of the navigation system's attitude error angle in the X, Y, and Z axes, respectively. For carrier height error, It contains three accelerometer zero bias errors in sequence. and the corresponding zero bias error of the three gyroscopes , This includes, in sequence, the projections of the accelerometers on the non-rotational axes onto each horizontal axis. , , Let X be the components of the displacement angular velocity in the X, Y, and Z directions under the navigation system. , The components of the Earth's rotational angular velocity in the Y and Z directions in the navigation system. Let be the radius of curvature of the meridian. The radius of curvature of the circle is denoted as . Latitude For height, , , Let X represent the components of velocity in the X, Y, and Z directions in the navigation system. , , These are the specific force values ​​measured by three accelerometers in the navigation system. Representing vectors Antisymmetric matrix, subscript This represents the value from the previous sampling period. Indicates the sampling interval; The measurement equation is established based on the second formula, which includes: ， ， ， in, The latitude is the coordinate for inertial navigation system calculation. For the longitude calculated by the inertial navigation system, The altitude calculated by the inertial navigation system. The latitude of the local location The longitude of the local location The altitude of the location For measuring noise; Kalman filtering is performed based on the state equation and the measurement equation to correct inertial navigation system parameter errors and device parameter errors in real time.

9. A self-aligning device for a rotating inertial navigation system that implements the method of claim 1, characterized in that, It includes, The coarse alignment module is used to obtain the orientation of the high-precision gyroscope in the initial strapdown position using a fast coarse alignment method, and to rotate the high-precision gyroscope to a set angle in the east-west direction according to the orientation of the high-precision gyroscope. A rotation module is used to obtain the initial attitude of the first position with the orientation of the high-precision gyroscope after rotation as the first position, and rotate the high-precision gyroscope to a second position symmetrical to the first position with the celestial axis as the central axis. A fine alignment module is used to perform fine alignment by using filtering estimation during the rotation from the first position to the second position and to obtain the attitude at the end of the rotation at the second position. The navigation tracking module is used to perform navigation tracking and complete self-alignment based on the attitude at the end of the second position and the inertial navigation data during the process of rotating around the celestial axis to the initial strapdown position.

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