A self-calibration method for missile-borne inertial system based on azimuth transfer mode

Abstract

The present invention relates to a missile-borne inertial system self-calibration method based on an azimuth transfer mode, belonging to the technical field of navigation, guidance and control. After a missile is released from a warehouse, the present invention uses optical aiming to transfer the initial azimuth to the missile-borne inertial system, 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, judge the error observable accuracy, and realize redundant self-calibration based on azimuth reference transfer. 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 system. 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.

Description

Technical Field

[0001] The present 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 a missile is released from storage, high-precision equipment is generally used to set the heading of the inertial measurement device to reduce the impact of heading errors. There are mainly two methods: the first is to obtain the azimuth reference from astronomical measurements and set it to the inertial navigation system, which is called optical aiming. The accuracy of optical aiming can reach the arc second level. This operation can be completed before leaving the storage and before launch. 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 storage and while on the move. Both methods essentially perform azimuth correction on the inertial measurement device. The present invention is based on the optical aiming azimuth reference setting before moving, and further observes the inertial navigation calibration error term, thereby improving its self-calibration accuracy. 1. On-the-move state: The missile-borne inertial navigation system is powered on once on the launch vehicle, and after completing the azimuth setting for leaving the storage, it is directly transported to the launch site. During this period, the missile-borne inertial navigation system is powered on and directly performs navigation. 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. Ballistic missiles are generally launched vertically, so erection operations are 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 system self-calibration method based on the azimuth transfer mode. After the missile is released from the warehouse, the present invention uses optical aiming to transfer the initial azimuth to the missile-borne inertial system, establishes an azimuth reference error transfer equation, improves the observability of the fiber optic gyroscope zero-position error through zero-speed and azimuth matching, and uses a Kalman filter to estimate the fiber optic gyroscope zero-position error to determine the error observability accuracy, thereby achieving redundant self-calibration based on azimuth reference transfer.

[0004] The present invention provides a missile-borne inertial system 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, expressed as follows:

[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 Sky Navigation Coordinate System i with reference to the Northeast Sky Navigation Coordinate System i.

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

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

[0010] Step S3: Obtain the roll angle and pitch angle of the missile-borne inertia group; establish a static base strapdown matrix based on the roll angle, pitch angle and azimuth angle of the missile-borne inertia group, and 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 calculation 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 the 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 missile-borne inertial group velocity error in the northeastern sky navigation coordinate system based on the simplified attitude error and the equivalent fiber optic gyroscope zero position error in the northeastern 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 missile-borne inertial group velocity error in the northeast sky navigation coordinate system is:

[0021]

[0022] in, is the acceleration error matrix of the missile-borne inertial group in the northeast sky navigation coordinate system, is the eastward acceleration error of the missile-borne inertial group in the northeast sky navigation coordinate system, is the north acceleration error of the missile-borne inertial group in the northeast sky navigation coordinate system, is the vertical acceleration error of the missile-borne inertial group in the northeast sky navigation coordinate system, Indicates the angular velocity error of the nth accelerometer in the vertical direction in the northeast sky navigation coordinate system, Indicates the angular velocity error of the nth accelerometer in the north direction in the northeast sky navigation coordinate system, g represents the acceleration of gravity, 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: Obtain the relationship between the attitude error and the velocity differential equation, which is expressed as:

[0024]

[0025] in, is the eastward velocity error of the missile-borne inertial group 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 vertical angular velocity of the fiber optic gyroscope in the northeast sky navigation coordinate system, ω N It represents the north angular velocity of the fiber optic gyroscope in the northeast sky navigation coordinate system.

[0026] Under the observation condition of a static base, the relationship between the velocity observation quantity 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 terms in the static state are obtained;

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

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

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

[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 Indicates the angular velocity error in the vertical direction in the northeast sky navigation coordinate system;

[0034] aB E ω U / gω N Represents the coupling term of acceleration zero error and gravity acceleration, aB E is the eastward acceleration observed 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: The zero position error of the equivalent east-pointing fiber optic gyroscope in the northeast sky navigation coordinate system is expressed as a heading angle error, which is expressed as follows:

[0036]

[0037] Among them, ω N is the angular velocity of the fiber optic gyroscope 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 gyroscope. It is usually used to describe the error value caused by the zero position deviation of the gyroscope in the heading angle measurement. gB E 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 gyroscope under the optical aiming reference;

[0039] Step S11: using each fiber optic gyroscope in the missile-borne inertial group as a target fiber optic gyroscope, and performing three-axis configuration on each fiber optic gyroscope to obtain multiple fiber optic gyroscope configuration schemes;

[0040] Performing scenario excitation on each of the fiber optic gyroscope configuration schemes, and obtaining the optimal configuration scheme based on the observability optimal allocation self-calibration method and the fiber optic gyroscope zero position error allocation weight under the 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 using the optimal configuration solution in step S11 include:

[0043] Taking each fiber optic gyroscope target 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] Add scenario excitation to each FOG three-axis configuration scheme to obtain the null factor error of each target FOG;

[0045] Obtaining the convergence of the scale factor error of the target fiber optic gyroscope in each three-axis configuration scheme through a Kalman filter; selecting the three-axis configuration scheme 1 corresponding to the target fiber optic gyroscope corresponding to the scale factor error with the largest convergence 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 optical aiming 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 utilizes 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 purposes of illustrating particular embodiments and are not to be considered limiting of the invention.

[0052] Figure 1Schematic diagram of the convergence curve of the static fiber optic gyroscope zero position error covariance after azimuth reference transfer in an embodiment of the present invention;

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

[0054] In order to more clearly understand the above-mentioned objects, features and advantages of the present invention, the present invention is further described in detail below with reference to 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. Therefore, the scope of protection 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 system calibration method based on the azimuth transfer mode. In order to illustrate the effectiveness of the method proposed by the present invention, the above technical solution of the present invention is described in detail below through a specific embodiment. 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, expressed as follows:

[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 Sky Navigation Coordinate System i with reference to the Northeast Sky Navigation Coordinate System i.

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

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

[0061] Step S3: Obtain the roll angle and pitch angle of the missile-borne inertia group; establish a static base strapdown matrix based on the roll angle, pitch angle and azimuth angle of the missile-borne inertia group, and 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 calculation 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 the 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 missile-borne inertial group velocity error in the northeastern sky navigation coordinate system based on the simplified attitude error and the equivalent fiber optic gyroscope zero position error in the northeastern 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 missile-borne inertial group velocity error in the northeast sky navigation coordinate system is:

[0072]

[0073] in, is the acceleration error matrix of the missile-borne inertial group in the northeast sky navigation coordinate system, is the eastward acceleration error of the missile-borne inertial group in the northeast sky navigation coordinate system, is the north acceleration error of the missile-borne inertial group in the northeast sky navigation coordinate system, is the vertical acceleration error of the missile-borne inertial group in the northeast sky navigation coordinate system, Indicates the angular velocity error of the nth accelerometer in the vertical direction in the northeast sky navigation coordinate system, Indicates the angular velocity error of the nth accelerometer in the north direction in the northeast sky navigation coordinate system, g represents the acceleration of gravity, 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: Obtain the relationship between the attitude error and the velocity differential equation, which is expressed as:

[0075]

[0076] in, is the eastward velocity error of the missile-borne inertial group 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 vertical angular velocity of the fiber optic gyroscope in the northeast sky navigation coordinate system, ω N It represents the north angular velocity of the fiber optic gyroscope in the northeast sky navigation coordinate system.

[0077] Under the observation condition of a static base, the relationship between the velocity observation quantity 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 terms in the static state are obtained;

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

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

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

[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 Indicates the angular velocity error in the vertical direction in the northeast sky navigation coordinate system;

[0085] aB E ω U / gω N Represents the coupling term of acceleration zero error and gravity acceleration, aB E is the eastward acceleration observed 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: The zero position error of the equivalent east-pointing fiber optic gyroscope in the northeast sky navigation coordinate system is expressed as a heading angle error, which is expressed as follows:

[0087]

[0088] Among them, ω N is the angular velocity of the fiber optic gyroscope 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 gyroscope. It is usually used to describe the error value caused by the zero position deviation of the gyroscope in the heading angle measurement. gB E 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 gyroscope under the optical aiming reference;

[0090] Step S11: using each fiber optic gyroscope in the missile-borne inertial group as a target fiber optic gyroscope, and performing three-axis configuration on each fiber optic gyroscope to obtain multiple fiber optic gyroscope configuration schemes;

[0091] Performing scenario excitation on each of the fiber optic gyroscope configuration schemes, and obtaining the optimal configuration scheme based on the observability optimal allocation self-calibration method and the fiber optic gyroscope zero position error allocation weight under the 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 using the optimal configuration solution in step S11 include:

[0094] Taking each fiber optic gyroscope target 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] Add scenario excitation to each FOG three-axis configuration scheme to obtain the null factor error of each target FOG;

[0096] Obtaining the convergence of the scale factor error of the target fiber optic gyroscope in each three-axis configuration scheme through a Kalman filter; selecting the three-axis configuration scheme 1 corresponding to the target fiber optic gyroscope corresponding to the scale factor error with the largest convergence 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 distribution table of fiber optic gyroscope zero position error 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, some coordinate systems have different construction values ​​and only have five weight distribution methods when allocating errors, especially when there is an optical aiming reference or degree of freedom restrictions.

[0101] As shown in the table above, based on azimuth and zero-velocity observations, the observability of the FOG a-fog and b-fog zero-position errors increases to approximately 50%. The corresponding observability of the c-fog and d-fog zero-position errors increases to approximately 62%, and the observability of the e-fog and f-fog zero-position errors increases to 31.1%. Clearly, the azimuth reference binding has the greatest effect on the c-fog and d-fog zero-position errors; it has a smaller effect on the e-fog and f-fog zero-position errors; and it has no effect on other FOG and accelerometer error terms.

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

Claims

1. A missile-borne inertial system calibration method based on the 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 sighting system; The azimuth angle of the missile-borne inertial group is obtained based on the northeastern sky navigation coordinate system of the optical aiming system; Step S2: Apply scene excitation to the missile-borne inertia group; Based on the azimuth baseline of the missile-borne inertial system, the zero position error and attitude error of each fiber optic gyroscope in the missile-borne inertial system 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 calculation 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 missile-borne inertial group velocity error in the northeastern sky navigation coordinate system based on the simplified attitude error and the equivalent fiber optic gyroscope zero position error in the northeastern 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 stationary state based on the simplified attitude error; Based on the heading angle error related terms in the static state, the coupling terms of acceleration zero error and gravity acceleration are established; Based on the coupling term of acceleration zero error and gravity acceleration, the zero error of the equivalent east-pointing fiber optic gyroscope in the northeast sky navigation coordinate system is obtained; Step S9: characterize 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 gyroscope under the optical aiming reference; Step S11: Calibrate the heading angle error of the fiber optic gyroscope calibration model based on the observability optimal distribution self-calibration method and the fiber optic gyroscope zero position error distribution weight under the optical aiming reference.

2. The missile-borne inertial system 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 inertia group.

3. The missile-borne inertial system 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 gyroscope zero position error in the northeast sky navigation coordinate system.

4. The missile-borne inertial system calibration method based on the azimuth transfer mode according to claim 1 is characterized in that: The acceleration zero 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 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 Indicates the angular velocity error in the vertical direction in the northeast sky navigation coordinate system; aBE ωU / gω N Represents the coupling term of acceleration zero error and gravity acceleration, aB E is the eastward acceleration observed 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 system 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: Obtain the relationship between attitude error and velocity differential equation based on the simplified attitude error; Obtain the relationship between the velocity observation and its corresponding derivative under the observation condition of a stationary base; 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 terms in the static state are obtained; Based on the heading angle error related terms in the static state, the coupling terms of acceleration zero error and gravity acceleration are established; Based on the coupling term of acceleration zero position error and gravity acceleration, the zero position error of the equivalent east-pointing fiber optic gyroscope in the northeast sky navigation coordinate system is obtained.

6. The missile-borne inertial system 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 gyroscope 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 gyroscope. It is usually used to describe the error value caused by the zero position deviation of the gyroscope in the heading angle measurement. gB E is the zero position error in the east direction in the northeast sky navigation coordinate system.

7. The missile-borne inertial system 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 gyroscope 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 the optimal configuration scheme based on the observability optimal allocation self-calibration method and the fiber optic gyroscope zero position error allocation weight under the 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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