Missile-borne inertial measurement unit self-calibration method based on azimuth transfer mode

Abstract

The invention relates to a missile-borne inertial measurement unit self-calibration method based on an orientation transfer mode, and belongs to the technical field of navigation guidance and control. After a missile is delivered out of a warehouse, an initial orientation is transferred to a missile-borne inertial measurement unit in an optical aiming mode, an orientation reference error transfer equation is established, and the self-calibration of the missile-borne inertial measurement unit is realized in a zero-speed and orientation matching mode; the observability of the zero error of the fiber-optic gyroscope is improved, the zero error of the fiber-optic gyroscope is estimated by using a Kalman filter, the observable precision of the error is judged, and the transmission redundancy self-calibration based on the azimuth reference is realized; according to the method, the high-precision light aiming system is used as a truth value reference, so that the attitude precision of the missile-borne inertial measurement unit is improved; the self-calibration capability of a missile system is improved, and the navigation precision and reliability of a missile in a complex environment are enhanced.

Description

Technical Field

[0001] The invention relates to the field of navigation, guidance and control technology, and in particular to a missile-borne inertial group self-calibration method based on an azimuth transfer mode. Background Art

[0002] Before the missile is released from the warehouse, high-precision equipment is generally used to bind the heading of the inertial measurement device to reduce the impact of the heading error. There are mainly two ways: the first is to obtain it by astronomical measurement, and bind the azimuth reference to the inertial navigation, that is, optical aiming. The optical aiming accuracy can reach the arc second level. This operation can be completed before leaving the warehouse and before launching; the second is to measure through the high-precision positioning and orientation system of the launch vehicle, and directly transfer the azimuth reference. This process can be completed after leaving the warehouse and during travel; both are essentially azimuth correction of the inertial measurement device. The present invention is based on the optical aiming azimuth reference binding before traveling, and further observes the inertial navigation calibration error term, thereby improving its self-calibration accuracy. 1. Traveling state: The missile-borne inertial navigation is powered on once on the launch vehicle, and after completing the out-of-warehouse azimuth binding, it is directly transported to the launch position. During this period, the missile-borne inertial navigation is powered on and directly navigates. 2. Standby position state: After the launch vehicle enters the standby position, other operations are generally not allowed. It only stays at the standby position for two to three minutes and then launches. For ballistic missiles, they are generally launched vertically, so erection operation is performed. The present invention utilizes the erection maneuver before launch to further calibrate the inertial guidance error term, thereby improving the self-calibration accuracy. Summary of the invention

[0003] In view of the above problems, the present invention provides a missile-borne inertial group self-calibration method based on an azimuth transfer mode. After the missile is released from the warehouse, the present invention uses an optical aiming method to transfer the initial azimuth to the missile-borne inertial group, establishes an azimuth reference error transfer equation, improves the observability of the fiber optic gyroscope zero position error through a zero speed and azimuth matching method, and uses a Kalman filter to estimate the fiber optic gyroscope zero position error, judges the error observable accuracy, and realizes redundant self-calibration based on azimuth reference transfer;

[0004] The present invention provides a missile-borne inertial group calibration method based on an azimuth transfer mode, comprising:

[0005] Step S1, determining the body coordinates of the missile-borne inertial group and the northeast sky navigation coordinate system of the optical sighting system; obtaining the azimuth of the missile-borne inertial group based on the northeast sky navigation coordinate system of the optical sighting system, the expression is;

[0006]

[0007] in, is the output of the angular velocity of the body coordinate system b relative to the northeast sky navigation coordinate system i in the body coordinate system b, is the rotation matrix from the northeast sky navigation coordinate system i to the local coordinate system b, It is the output of the earth's rotation angular velocity in the northeast celestial navigation coordinate system i with the northeast celestial navigation coordinate system i as the reference.

[0008] Step S2, applying scene excitation to the missile-borne inertial group; obtaining the zero position error and attitude error of each fiber optic gyroscope in the missile-borne inertial group based on the azimuth baseline of the missile-borne inertial group;

[0009] Preferably, the scene excitation includes the movement of the missile-borne inertial unit;

[0010] Step S3, obtaining the roll angle and pitch angle of the missile-borne inertia group; establishing a static base strapdown matrix based on the roll angle, pitch angle and azimuth angle of the missile-borne inertia group, the expression is:

[0011]

[0012] Where, T is the stationary base strapdown matrix, γ is the roll angle, θ is the pitch angle, and A is the optical aiming azimuth.

[0013] Step S4, substituting the zero position error of each fiber optic gyroscope into the stationary base strapdown matrix to obtain the equivalent fiber optic gyroscope zero position error in the northeast sky navigation coordinate system;

[0014] Step S5, obtaining a velocity observation value of the missile-borne inertial system in the navigation coordinate system; performing differential derivation on the velocity observation value to obtain a processed velocity observation value; obtaining an attitude error of the missile-borne inertial system based on the processed velocity observation value and an equivalent fiber optic gyroscope zero position error in the northeast sky navigation coordinate system;

[0015] Step S6, simplify the attitude error of the missile-borne inertia group to obtain a simplified attitude error, which is expressed as:

[0016]

[0017] in, is the simplified attitude error, φ is the carrier attitude, gB n is the equivalent fiber optic gyro zero position error in the northeast sky navigation coordinate system;

[0018] Step S7, obtaining the velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system based on the simplified attitude error and the equivalent fiber optic gyroscope zero position error in the northeast sky navigation coordinate system;

[0019] Calibrate the velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system;

[0020] Preferably, the expression of the velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system is:

[0021]

[0022] in, is the acceleration error matrix of the missile-borne inertial unit in the northeast sky navigation coordinate system, is the eastward acceleration error of the missile-borne inertial unit in the northeast sky navigation coordinate system, is the north acceleration error of the missile-borne inertial unit in the northeast sky navigation coordinate system, is the vertical acceleration error of the missile-borne inertial unit in the northeast sky navigation coordinate system, It represents the angular velocity error of the nth accelerometer in the vertical direction in the northeast sky navigation coordinate system. It represents the angular velocity error of the nth accelerometer in the north direction in the northeast sky navigation coordinate system, g represents the gravitational acceleration, φ E It represents the value of the attitude angle of the missile-borne inertial group corresponding to the east direction in the northeast sky navigation coordinate system, φ N It represents the value of the attitude angle of the missile-borne inertial group corresponding to the north direction in the northeast sky navigation coordinate system, φ U It indicates the value of the attitude angle of the missile-borne inertial group corresponding to the vertical direction in the northeast sky navigation coordinate system, the attitude angle of the missile-borne inertial group, such as pitch, yaw and roll angle.

[0023] Step S8, obtaining the relationship between the attitude error and the velocity differential equation, the expression is:

[0024]

[0025] in, is the eastward velocity error of the missile-borne inertial unit in the northeast sky navigation coordinate system, is the north velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system, aB E is the eastward zero position error of the accelerometer in the northeast sky navigation coordinate system, aB N is the north zero position error in the northeast sky navigation coordinate system, gB U is the vertical installation error in the northeast sky navigation coordinate system, ω U represents the angular velocity of the fiber optic gyroscope in the vertical direction in the northeast sky navigation coordinate system, ω N It represents the angular velocity of the fiber optic gyroscope in the north direction in the northeast sky navigation coordinate system.

[0026] Under the condition of static base observation, the relationship between the velocity observation and its corresponding derivative is obtained as follows:

[0027]

[0028] Based on the relationship between attitude error and velocity differential equation, and the relationship between velocity observation and its corresponding derivative, the heading angle error related items in the static state are obtained;

[0029] Based on the heading angle error related items in the static state, the coupling items of the acceleration zero position error and the gravity acceleration are established;

[0030] Based on the coupling term of acceleration zero position error and gravity acceleration, the zero position error of the equivalent eastward fiber optic gyroscope in the northeast sky navigation coordinate system is obtained;

[0031] Preferably, the acceleration zero position error and gravity acceleration coupling term are expressed as:

[0032] aB E ω U / gω N =(aB E / g)·(ω U / ω N )=tan L·aB E / g

[0033] Among them, aB E ω U It represents the coupling term between the eastward acceleration observation and the vertical angular velocity in the northeast sky navigation coordinate system, gω N It represents the angular velocity error in the vertical direction in the northeast sky navigation coordinate system;

[0034] aB E ω U / gω N represents the acceleration zero position error and gravity acceleration coupling term, aB E is the observed eastward acceleration in the northeast sky navigation coordinate system, ω U represents the angular velocity of the fiber optic gyroscope in the vertical direction in the northeast sky navigation coordinate system, ω N It represents the angular velocity of the fiber optic gyroscope in the north direction in the northeast sky navigation coordinate system, and L is the local latitude.

[0035] Step S9: characterize the zero position error of the equivalent east-direction fiber optic gyroscope in the northeast sky navigation coordinate system as a heading angle error, and the expression is:

[0036]

[0037] Among them, ω N is the angular velocity of the fiber optic gyro in the north direction in the northeast sky navigation coordinate system, φ U The heading angle error is caused by the zero position error of the fiber optic gyro. It is usually used to describe the error value caused by the zero position deviation of the gyro in the heading angle measurement. gB E It is the zero position error in the east direction in the northeast sky navigation coordinate system.

[0038] Step S10, setting the zero position error distribution weight of the fiber optic gyro under the optical aiming reference;

[0039] Step S11, taking each fiber optic gyroscope in the missile-borne inertial group as a target fiber optic gyroscope, and performing three-axis configuration respectively, to obtain a plurality of fiber optic gyroscope configuration schemes;

[0040] Performing scenario excitation on each of the fiber optic gyroscope configuration schemes, and obtaining an optimal configuration scheme based on an observability optimal allocation self-calibration method and a fiber optic gyroscope zero position error allocation weight under an optical aiming reference;

[0041] The heading angle error of the fiber optic gyroscope calibration model is calibrated using the optimal configuration scheme.

[0042] Preferably, the specific steps of calibrating the heading angle of the fiber optic gyroscope calibration model by using the optimal configuration scheme in step S11 include:

[0043] Taking each fiber optic gyroscope mark as a target fiber optic gyroscope, establishing a corresponding fiber optic gyroscope three-axis configuration scheme, and obtaining multiple fiber optic gyroscope three-axis configuration schemes;

[0044] Adding scene excitation to each fiber optic gyroscope three-axis configuration scheme, obtaining the zero factor error of each target fiber optic gyroscope;

[0045] The convergence of the scale factor error of the target fiber optic gyroscope in each three-axis configuration scheme is obtained by using a Kalman filter; the three-axis configuration scheme 1 corresponding to the target fiber optic gyroscope corresponding to the scale factor error with the largest convergence degree is selected as the optimal configuration scheme 1;

[0046] The optimal configuration scheme is used to calibrate the heading angle error of the determined fiber optic gyroscope calibration model.

[0047] Compared with the prior art, the present invention has at least the following beneficial effects:

[0048] (1) The present invention transmits the initial azimuth to the missile-borne inertial group through an optical aiming method, and adopts a zero-speed and azimuth matching method to improve the observability of the zero-position error of the fiber optic gyroscope;

[0049] (2) The present invention uses a high-precision optical aiming system as a true value reference to improve the attitude accuracy of the missile-borne inertial group, and estimates the error through a Kalman filter to improve the observation accuracy;

[0050] (3) The present invention improves the self-calibration capability of the missile system and enhances the navigation accuracy and reliability of the missile in complex environments. BRIEF DESCRIPTION OF THE DRAWINGS

[0051] The drawings are only for the purpose of illustrating particular embodiments and are not to be construed as limiting the invention.

[0052] Figure 1It is a schematic diagram of a convergence curve of the static fiber optic gyroscope zero position error covariance after the azimuth reference transmission in an embodiment of the present invention;

[0053] Figure 2 It is a schematic diagram of the principle diagram of the zero bias error covariance convergence curve of the east accelerometer in an embodiment of the present invention. DETAILED DESCRIPTION

[0054] In order to more clearly understand the above-mentioned purpose, features and advantages of the present invention, the present invention is further described in detail below in conjunction with the accompanying drawings and specific embodiments. It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments can be combined with each other. In addition, the present invention can also be implemented in other ways different from those described herein, and therefore, the protection scope of the present invention is not limited by the specific embodiments disclosed below.

[0055] A specific embodiment of the present invention, as Figure 1-2 , discloses a missile-borne inertial group calibration method based on the azimuth transfer mode. In order to illustrate the effectiveness of the method proposed in the present invention, the above technical solution of the present invention is described in detail through a specific embodiment below. The specific implementation steps are as follows:

[0056] Step S1, determining the body coordinates of the missile-borne inertial group and the northeast sky navigation coordinate system of the optical sighting system; obtaining the azimuth of the missile-borne inertial group based on the northeast sky navigation coordinate system of the optical sighting system, the expression is;

[0057]

[0058] in, is the output of the angular velocity of the body coordinate system b relative to the northeast sky navigation coordinate system i in the body coordinate system b, is the rotation matrix from the northeast sky navigation coordinate system i to the local coordinate system b, It is the output of the earth's rotation angular velocity in the northeast celestial navigation coordinate system i with the northeast celestial navigation coordinate system i as the reference.

[0059] Step S2, applying scene excitation to the missile-borne inertial group; obtaining the zero position error and attitude error of each fiber optic gyroscope in the missile-borne inertial group based on the azimuth baseline of the missile-borne inertial group;

[0060] Preferably, the scene excitation includes the movement of the missile-borne inertial unit;

[0061] Step S3, obtaining the roll angle and pitch angle of the missile-borne inertia group; establishing a static base strapdown matrix based on the roll angle, pitch angle and azimuth angle of the missile-borne inertia group, the expression is:

[0062]

[0063] Where, T is the stationary base strapdown matrix, γ is the roll angle, θ is the pitch angle, and A is the optical aiming azimuth.

[0064] Step S4, substituting the zero position error of each fiber optic gyroscope into the stationary base strapdown matrix to obtain the equivalent fiber optic gyroscope zero position error in the northeast sky navigation coordinate system;

[0065] Step S5, obtaining a velocity observation value of the missile-borne inertial system in the navigation coordinate system; performing differential derivation on the velocity observation value to obtain a processed velocity observation value; obtaining an attitude error of the missile-borne inertial system based on the processed velocity observation value and an equivalent fiber optic gyroscope zero position error in the northeast sky navigation coordinate system;

[0066] Step S6, simplify the attitude error of the missile-borne inertia group to obtain a simplified attitude error, which is expressed as:

[0067]

[0068] in, is the simplified attitude error, φ is the carrier attitude, gB n is the equivalent fiber optic gyro zero position error in the northeast sky navigation coordinate system;

[0069] Step S7, obtaining the velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system based on the simplified attitude error and the equivalent fiber optic gyroscope zero position error in the northeast sky navigation coordinate system;

[0070] Calibrate the velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system;

[0071] Preferably, the expression of the velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system is:

[0072]

[0073] in, is the acceleration error matrix of the missile-borne inertial unit in the northeast sky navigation coordinate system, is the eastward acceleration error of the missile-borne inertial unit in the northeast sky navigation coordinate system, is the north acceleration error of the missile-borne inertial unit in the northeast sky navigation coordinate system, is the vertical acceleration error of the missile-borne inertial unit in the northeast sky navigation coordinate system, It represents the angular velocity error of the nth accelerometer in the vertical direction in the northeast sky navigation coordinate system. It represents the angular velocity error of the nth accelerometer in the north direction in the northeast sky navigation coordinate system, g represents the gravitational acceleration, φ E It represents the value of the attitude angle of the missile-borne inertial group corresponding to the east direction in the northeast sky navigation coordinate system, φ NIt represents the value of the attitude angle of the missile-borne inertial group corresponding to the north direction in the northeast sky navigation coordinate system, φ U It indicates the value of the attitude angle of the missile-borne inertial group corresponding to the vertical direction in the northeast sky navigation coordinate system, the attitude angle of the missile-borne inertial group, such as pitch, yaw and roll angle.

[0074] Step S8, obtaining the relationship between the attitude error and the velocity differential equation, the expression is:

[0075]

[0076] in, is the eastward velocity error of the missile-borne inertial unit in the northeast sky navigation coordinate system, is the north velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system, aB E is the eastward zero position error of the accelerometer in the northeast sky navigation coordinate system, aB N is the north zero position error in the northeast sky navigation coordinate system, gB U is the vertical installation error in the northeast sky navigation coordinate system, ω U represents the angular velocity of the fiber optic gyroscope in the vertical direction in the northeast sky navigation coordinate system, ω N It represents the angular velocity of the fiber optic gyroscope in the north direction in the northeast sky navigation coordinate system.

[0077] Under the condition of static base observation, the relationship between the velocity observation and its corresponding derivative is obtained as follows:

[0078]

[0079] Based on the relationship between attitude error and velocity differential equation, and the relationship between velocity observation and its corresponding derivative, the heading angle error related items in the static state are obtained;

[0080] Based on the heading angle error related items in the static state, the coupling items of the acceleration zero position error and the gravity acceleration are established;

[0081] Based on the coupling term of acceleration zero position error and gravity acceleration, the zero position error of the equivalent eastward fiber optic gyroscope in the northeast sky navigation coordinate system is obtained;

[0082] Preferably, the acceleration zero position error and gravity acceleration coupling term are expressed as:

[0083] aB E ω U / gω N =(aB E / g)·(ω U / ω N )=tan L·aB E / g

[0084] Among them, aB E ω U It represents the coupling term between the eastward acceleration observation and the vertical angular velocity in the northeast sky navigation coordinate system, gω N It represents the angular velocity error in the vertical direction in the northeast sky navigation coordinate system;

[0085] aB E ω U / gω N represents the acceleration zero position error and gravity acceleration coupling term, aB E is the observed eastward acceleration in the northeast sky navigation coordinate system, ω U represents the angular velocity of the fiber optic gyroscope in the vertical direction in the northeast sky navigation coordinate system, ω N It represents the angular velocity of the fiber optic gyroscope in the north direction in the northeast sky navigation coordinate system, and L is the local latitude.

[0086] Step S9: characterize the zero position error of the equivalent east-direction fiber optic gyroscope in the northeast sky navigation coordinate system as a heading angle error, and the expression is:

[0087]

[0088] Among them, ω N is the angular velocity of the fiber optic gyro in the north direction in the northeast sky navigation coordinate system, φ U The heading angle error is caused by the zero position error of the fiber optic gyro. It is usually used to describe the error value caused by the zero position deviation of the gyro in the heading angle measurement. gB E It is the zero position error in the east direction in the northeast sky navigation coordinate system.

[0089] Step S10, setting the zero position error distribution weight of the fiber optic gyro under the optical aiming reference;

[0090] Step S11, taking each fiber optic gyroscope in the missile-borne inertial group as a target fiber optic gyroscope, and performing three-axis configuration respectively, to obtain a plurality of fiber optic gyroscope configuration schemes;

[0091] Performing scenario excitation on each of the fiber optic gyroscope configuration schemes, and obtaining an optimal configuration scheme based on an observability optimal allocation self-calibration method and a fiber optic gyroscope zero position error allocation weight under an optical aiming reference;

[0092] The heading angle error of the fiber optic gyroscope calibration model is calibrated using the optimal configuration scheme.

[0093] Preferably, the specific steps of calibrating the heading angle of the fiber optic gyroscope calibration model by using the optimal configuration scheme in step S11 include:

[0094] Taking each fiber optic gyroscope mark as a target fiber optic gyroscope, establishing a corresponding fiber optic gyroscope three-axis configuration scheme, and obtaining multiple fiber optic gyroscope three-axis configuration schemes;

[0095] Adding scene excitation to each fiber optic gyroscope three-axis configuration scheme, obtaining the zero factor error of each target fiber optic gyroscope;

[0096] The convergence of the scale factor error of the target fiber optic gyroscope in each three-axis configuration scheme is obtained by using a Kalman filter; the three-axis configuration scheme 1 corresponding to the target fiber optic gyroscope corresponding to the scale factor error with the largest convergence degree is selected as the optimal configuration scheme 1;

[0097] The optimal configuration scheme is used to calibrate the heading angle error of the determined fiber optic gyroscope calibration model.

[0098] Table 1 Weight table of zero position error distribution of fiber optic gyro under optical aiming reference

[0099]

[0100] As shown in Table 1, twenty coordinate systems are constructed based on the configuration structure of the regular dodecahedron. Among them, the construction values ​​of some coordinate systems are different, and there are only five weight distribution methods when allocating errors, especially under conditions with optical aiming datum or degree of freedom restrictions.

[0101] As can be seen from the table above, based on azimuth and zero-speed observations, the observability of the zero-position errors of the fiber optic gyroscope a-fog and b-fog is increased to about 50%; the corresponding observability of the zero-position errors of c-fog and d-fog is increased to about 62%; the observability of the zero-position errors of e-fog and f-fog is increased to 31.1%. Obviously, the binding of the azimuth reference has the greatest improvement on the zero-position errors of c-fog and d-fog; the improvement on the zero-position errors of e-fog and f-fog is small; and there is no improvement on other error items of the fiber optic gyroscope and accelerometer.

[0102] The above description is only a preferred specific implementation manner of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by any technician familiar with the technical field within the technical scope disclosed by the present invention should be covered within the protection scope of the present invention.

Claims

1. A missile-borne inertial unit calibration method based on azimuth transfer mode, characterized in that: include: Step S1, determining the body coordinates of the missile-borne inertial group and the northeast sky navigation coordinate system of the optical aiming system; The azimuth of the missile-borne inertial unit is obtained based on the northeastern sky navigation coordinate system of the optical aiming system; Step S2, applying scene excitation to the missile-borne inertia group; Based on the azimuth baseline of the missile-borne inertial group, the zero position error and attitude error of each fiber optic gyroscope in the missile-borne inertial group are obtained; Step S3, obtaining the roll angle and pitch angle of the missile-borne inertia group; establishing a static base strapdown matrix based on the roll angle, pitch angle and azimuth angle of the missile-borne inertia group; Step S4, substituting the zero position error of each fiber optic gyroscope into the stationary base strapdown matrix to obtain the equivalent fiber optic gyroscope zero position error in the northeast sky navigation coordinate system; Step S5, obtaining a velocity observation value of the missile-borne inertial system in the navigation coordinate system; performing differential derivation on the velocity observation value to obtain a processed velocity observation value; obtaining an attitude error of the missile-borne inertial system based on the processed velocity observation value and an equivalent fiber optic gyroscope zero position error in the northeast sky navigation coordinate system; Step S6, simplifying the attitude error of the missile-borne inertia group to obtain a simplified attitude error; Step S7, obtaining the velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system based on the simplified attitude error and the equivalent fiber optic gyroscope zero position error in the northeast sky navigation coordinate system; Calibrate the velocity error of the missile-borne inertial group in the northeast sky navigation coordinate system; Step S8, obtaining heading angle error related items in a static state based on the simplified attitude error; Based on the heading angle error related items in the static state, the zero position error of the equivalent eastward fiber optic gyroscope in the northeast sky navigation coordinate system is obtained; Step S9, characterizing the zero position error of the equivalent east-pointing fiber optic gyroscope in the northeast sky navigation coordinate system as a heading angle error; Step S10, setting the zero position error distribution weight of the fiber optic gyro under the optical aiming reference; Step S11, based on the observability optimal allocation self-calibration method and the fiber optic gyro zero position error allocation weight under the optical aiming reference, calibrate the heading angle error of the fiber optic gyro calibration model.

2. The missile-borne inertial unit calibration method based on the azimuth transfer mode according to claim 1 is characterized in that: The scene excitation in step S2 includes the movement of the missile-borne inertial group.

3. The missile-borne inertial unit calibration method based on the azimuth transfer mode according to claim 1 is characterized in that: The simplified expression of the posture error in step S6 is: in, is the simplified attitude error, φ is the carrier attitude, gB n is the equivalent fiber optic gyro zero position error in the northeast sky navigation coordinate system.

4. The missile-borne inertial unit calibration method based on the azimuth transfer mode according to claim 1 is characterized in that: The acceleration zero position error and gravity acceleration coupling term in step S8 are expressed as follows: aB E ω U / gω N =(aB E / g)·(ω U / ω N )=tanL·aB E / g Among them, aB E ωU represents the coupling term between the eastward acceleration observation and the vertical angular velocity in the northeast sky navigation coordinate system, gω N It represents the angular velocity error in the vertical direction in the northeast sky navigation coordinate system; aB E ω U / gω N represents the acceleration zero position error and gravity acceleration coupling term, aB E is the observed eastward acceleration in the northeast sky navigation coordinate system, ω U represents the angular velocity of the fiber optic gyroscope in the vertical direction in the northeast sky navigation coordinate system, ω N It represents the angular velocity of the fiber optic gyroscope in the north direction in the northeast sky navigation coordinate system, and L is the local latitude.

5. The missile-borne inertial unit calibration method based on the azimuth transfer mode according to claim 1 is characterized in that: The specific steps of obtaining the zero position error of the equivalent east-pointing fiber optic gyroscope in the northeast sky navigation coordinate system in step S8 include: The relationship between the attitude error and the velocity differential equation is obtained based on the simplified attitude error; Obtain the relationship between the velocity observation and its corresponding derivative under the condition of static base observation; Based on the relationship between attitude error and velocity differential equation, and the relationship between velocity observation and its corresponding derivative, the heading angle error related items in the static state are obtained; Based on the heading angle error related items in the static state, the coupling items of the acceleration zero position error and the gravity acceleration are established; Based on the coupling term of acceleration zero position error and gravity acceleration, the zero position error of the equivalent eastward fiber optic gyroscope in the northeast sky navigation coordinate system is obtained.

6. The missile-borne inertial unit calibration method based on the azimuth transfer mode according to claim 1 is characterized in that: The heading angle error in step S9 is expressed as: Among them, ω N is the angular velocity of the fiber optic gyro in the north direction in the northeast sky navigation coordinate system, φ U The heading angle error is caused by the zero position error of the fiber optic gyro. It is usually used to describe the error value caused by the zero position deviation of the gyro in the heading angle measurement. gB E It is the zero position error in the east direction in the northeast sky navigation coordinate system.

7. The missile-borne inertial unit calibration method based on the azimuth transfer mode according to claim 1 is characterized in that: The specific steps of calibrating the heading angle error of the fiber optic gyro calibration model in step S11 include: Each fiber optic gyroscope in the missile-borne inertial group is used as a target fiber optic gyroscope, and three-axis configurations are performed respectively to obtain multiple fiber optic gyroscope configuration schemes; Performing scenario excitation on each of the fiber optic gyroscope configuration schemes, and obtaining an optimal configuration scheme based on an observability optimal allocation self-calibration method and a fiber optic gyroscope zero position error allocation weight under an optical aiming reference; The heading angle error of the fiber optic gyroscope calibration model is calibrated using the optimal configuration scheme.

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