An airborne inertial navigation gyroscope zero bias estimation method combined with transmission alignment

Abstract

The application relates to an airborne inertial navigation gyro zero bias estimation method combined with transmission alignment, and belongs to the technical field of inertial navigation gyro zero bias estimation, and comprises the following steps: S1, constructing a relative installation error analytical expression model between a main inertial navigation system and a sub inertial navigation system; S2, obtaining a relative installation error based on the relative installation error analytical expression model constructed in step S1; S3, obtaining initial self-alignment error and self-alignment error after t time after alignment by using the relative installation error obtained in step S2; and S4, obtaining a gyro zero bias estimation result of the sub inertial navigation system based on the initial self-alignment error and the self-alignment error after t time after alignment obtained in step S3. The method provided by the application is high in reliability, strong in universality, high in operation efficiency of an algorithm, simple in operation, high in precision and good in practicability.

Description

Technical Field

[0001] This invention belongs to the field of inertial navigation gyroscope zero bias estimation technology, specifically involving an airborne inertial navigation gyroscope zero bias estimation method combined with transfer alignment. Background Technology

[0002] Although inertial navigation systems (INS) undergo calibration and compensation before leaving the factory, some system errors can change after prolonged storage, especially gyroscope bias. Therefore, gyroscope bias needs to be recalibrated during use. Furthermore, due to limitations in payload and cost, UAVs often carry low-precision INS, which acquire relatively low initial attitude accuracy through self-alignment. The presence of initial attitude errors and gyroscope bias leads to a rapid deterioration in the attitude and position accuracy of the UAV during flight. Therefore, without changing the hardware, the key to improving the accuracy of inertial navigation systems lies in enhancing their initial attitude accuracy and gyroscope bias compensation accuracy.

[0003] The proposed transfer alignment method effectively solves the problem of low initial attitude accuracy obtained by airborne inertial navigation system (INS) self-alignment. This method corrects the initial attitude of the airborne INS by estimating the relative installation angle between the high-precision INS (main INS) on the UAV launch platform and the airborne INS (sub-INS). However, few transfer alignment methods have implemented gyroscope zero-bias compensation based on the transfer alignment results to further improve the navigation performance of the inertial navigation system.

[0004] Existing methods for compensating gyroscope bias fall into two categories: laboratory calibration and online estimation. Laboratory calibration is limited by disassembly and is time-consuming, making it unsuitable for recalibrating inertial navigation systems after they leave the factory. Current online gyroscope bias calibration methods largely rely on other navigational information, such as GNSS, Doppler logs, and visual signals. GNSS fails in denied environments, Doppler logs are only suitable for unmanned surface vessels (USVs) and other maritime vessels, and visual signals are susceptible to light pollution. Therefore, a novel, robust gyroscope bias estimation method is needed to further improve the accuracy of airborne inertial navigation systems. Summary of the Invention

[0005] To address the problems existing in the prior art, the present invention aims to provide a method for estimating the zero bias of an airborne inertial navigation gyroscope in conjunction with transfer alignment. This method is highly reliable, versatile, efficient, simple to operate, highly accurate, and practical.

[0006] To achieve the above objectives, the present invention adopts the following technical solution: a method for estimating the zero bias of an airborne inertial navigation gyroscope combined with transfer alignment, comprising the following steps:

[0007] S1. Construct an analytical model of the relative installation error between the main inertial navigation system and the sub-inertial navigation system;

[0008] S2. Based on the analytical model of relative installation error constructed in step S1, obtain the relative installation error;

[0009] S3. Using the relative installation error obtained in step S2, obtain the initial self-alignment error and the self-alignment error after alignment time t.

[0010] S4. Based on the initial self-alignment error obtained in step S3 and the self-alignment error after alignment time t, the gyroscope zero-bias estimation result of the sub-inertial navigation system is obtained.

[0011] Furthermore, in step S1, the analytical model of the relative installation error between the main inertial navigation system and the sub-inertial navigation system is constructed as follows:

[0012]

[0013] In the formula, Main inertial navigation in t m-1 to t m The angular increment vector at time t.

[0014] Main inertial navigation in t m-1 to t m The angular increment of the X-gyroscope at time t, Main inertial navigation in t m-1 to t m The angular increment of the Y-gyroscope at time t. Main inertial navigation in t m-1 to t m The angular increment of the Z-gyroscope at time t;

[0015] antisymmetric matrix,

[0016]

[0017] The relative installation error vector between the master and sub-inertial navigation systems. Δβ 1,2 The pitch relative installation error of the master and slave inertial navigation systems, Δγ 1,2 The relative roll installation error of the master and sub-inertial navigation systems, Δα 1,2 The relative installation error of the heading of the main and sub-inertial navigation systems;

[0018] The relative attitude error vector caused by self-alignment error is determined by the following formula:

[0019]

[0020] In the formula, For t m The self-alignment error vector at time t.

[0021] δβ 1,2 (t m ) for t m The pitch relative attitude error at any given time, δγ 1，2 (t m ) for t m The relative roll attitude error at time δα 1，2 (t m ) for t m The relative attitude error of the heading at any given moment;

[0022] for antisymmetric matrix,

[0023]

[0024] For t m The relative attitude error vector at time t.

[0025] yes antisymmetric matrix,

[0026] Furthermore, in step S2, the specific method for obtaining the relative installation error based on the analytical model of relative installation error includes the following steps:

[0027] S201. Collect the attitude vector ψ1 output by the main inertial navigation system, the attitude vector ψ2 output by the sub-inertial navigation system, and the angular increment vector output by the main inertial navigation system within a time interval T.

[0028] Where s≥3, ψ1=[β1,γ1,α1] T ψ2=[β2,γ2,α2] T β1 is the pitch angle output by the main inertial navigation system, γ1 is the roll angle output by the main inertial navigation system, α1 is the yaw angle output by the main inertial navigation system, β2 is the pitch angle output by the sub-inertial navigation system, γ2 is the roll angle output by the sub-inertial navigation system, and α2 is the yaw angle output by the sub-inertial navigation system.

[0029] S202. Based on the attitude vector ψ1 output by the main inertial navigation system and the attitude vector ψ2 output by the sub-inertial navigation system obtained in step S201, use the formula... Obtain the relative attitude error vector Right now

[0030]

[0031] In the formula, β1(t) m ) for t m The pitch angle output by the master inertial navigation system at time t, γ1(t) m ) for t m The roll angle output by the master inertial navigation system at time t, α1(t) m ) for t m The heading angle output by the master inertial navigation system at time t, β2(t) m ) for t m The pitch angle output by the sub-inertial navigation system at time t, γ2(t) m ) for t m The roll angle output by the sub-inertial navigation system at time t, α2(t) m ) for t m The heading angle output by the sub-inertial navigation system at any given moment;

[0032] S203. Based on step S1, construct an analytical model of the relative installation error between the main inertial navigation system and the sub-inertial navigation system, and use the attitude vectors output by the main inertial navigation system, the attitude vectors output by the sub-inertial navigation system, the angular increment vector output by the main inertial navigation system, and the relative attitude error vector obtained in step S201, and solve for the relative installation error vector.

[0033] Specifically, the attitude vectors output by the main inertial navigation system (INS), the attitude vectors output by the sub-INS, the angular increment vector output by the main INS, and the relative attitude error vectors obtained in step S201 within a time period T are substituted into the analytical model of relative installation error to obtain s sets of equations:

[0034]

[0035] An equation has three variables: Δβ 1,2 ,Δγ 1,2 and Δα 1,2 Since s≥3 is guaranteed in step S201, the solution is obtained by the least squares method. The relative installation error vector can then be obtained.

[0036] Furthermore, in step S3, the specific method for obtaining the initial self-alignment error and the self-alignment error after alignment time t using the relative installation error obtained in step S2 includes the following steps:

[0037] S301. Collect the attitude vector ψ1(0) output at the initial moment after the main inertial navigation system is aligned and the attitude vector ψ2(0) output at the initial moment after the sub-inertial navigation system is aligned.

[0038] Where ψ1(0)=[β1(0),γ1(0),α1(0)] T,ψ2(0)=[β2(0),γ2(0),α2(0)] T β1(0) is the pitch angle output at the initial moment of the main inertial navigation system's alignment completion, γ1(0) is the roll angle output at the initial moment of the main inertial navigation system's alignment completion, α1(0) is the heading angle output at the initial moment of the main inertial navigation system's alignment completion, β2(0) is the pitch angle output at the initial moment of the sub-inertial navigation system's alignment completion, γ2(0) is the roll angle output at the initial moment of the sub-inertial navigation system's alignment completion, and α2(0) is the heading angle output at the initial moment of the sub-inertial navigation system's alignment completion.

[0039] S302. Based on the attitude vector ψ1(0) output at the initial moment of the main inertial navigation system alignment completion and the attitude vector ψ2(0) output at the initial moment of the sub-inertial navigation system alignment completion obtained in step S301, using the formula... Obtain the self-alignment error vector at the initial moment after alignment completion, i.e.

[0040]

[0041] S303. Collect the attitude vector ψ1(t) output after time t after the main inertial navigation system alignment is completed, and the attitude vector ψ2(t) output after time t after the sub-inertial navigation system alignment is completed.

[0042] Among them, ψ1(t)=[β1(t),γ1(t),α1(t)] T ,ψ2(t)=[β2(t),γ2(t),α2(t)] T β1(t) is the pitch angle output after time t after the main inertial navigation system (INS) alignment is completed, γ1(t) is the roll angle output after time t after the main INS alignment is completed, α1(t) is the heading angle output after time t after the main INS alignment is completed, β2(t) is the pitch angle output after time t after the sub-INS alignment is completed, γ2(t) is the roll angle output after time t after the sub-INS alignment is completed, and α2(t) is the heading angle output after time t after the sub-INS alignment is completed.

[0043] S304. Based on the attitude vector ψ1(t) output after time t following the completion of the main inertial navigation system alignment and the attitude vector ψ2(t) output after time t following the completion of the sub-inertial navigation system alignment, obtained in step S303, the formula is used... Obtain the self-alignment error vector at the initial moment after alignment completion, i.e.

[0044]

[0045] Furthermore, in step S4, the specific method for obtaining the gyroscope zero-bias estimation result of the sub-inertial navigation system based on the initial self-alignment error obtained in step S3 and the self-alignment error after alignment time t includes the following steps:

[0046] S401. Acquire the attitude matrix output at the initial moment after the sub-inertial navigation alignment is completed.

[0047] S402, Using the formula Obtain the gyroscope zero-bias estimation result for the sub-inertial navigation system; where, The gyroscope zero-bias estimation result is given for the sub-inertial navigation system, where ΔT is the sensor sampling time, and t≥100ΔT.

[0048] Compared with existing technologies, the present invention has the following advantages: The method provided by the present invention establishes an analytical model of the relative installation error between the main inertial navigation system (MSN) and the sub-MSN to ultimately obtain the relative attitude between the main and sub-MSNs. Based on the relative attitude between the main and sub-MSNs, the gyro zero bias of the airborne inertial navigation system is estimated online to further improve the navigation accuracy of the airborne inertial navigation system. It features high reliability, strong versatility, high algorithm efficiency, simple operation, high accuracy, and excellent practicality. The method provided by the present invention solves the problems of existing transfer alignment methods that do not perform gyro zero bias compensation based on the transfer alignment results to further improve the navigation performance of the inertial navigation system, and existing online gyro zero bias calibration methods that rely heavily on other navigation information. Attached Figure Description

[0049] Figure 1 This is a flowchart illustrating an airborne inertial navigation gyroscope zero-bias estimation method combined with transfer alignment according to the present invention. Detailed Implementation

[0050] The present invention will be further described below with reference to the accompanying drawings and specific embodiments, but the following embodiments are by no means intended to limit the present invention.

[0051] like Figure 1 As shown, a specific embodiment of the present invention provides a method for estimating the zero bias of an airborne inertial navigation gyroscope in conjunction with transfer alignment, comprising the following steps:

[0052] S1. Construct an analytical model of the relative installation error between the main inertial navigation system and the sub-inertial navigation system;

[0053] S2. Based on the analytical model of relative installation error constructed in step S1, obtain the relative installation error;

[0054] S3. Using the relative installation error obtained in step S2, obtain the initial self-alignment error and the self-alignment error after alignment time t.

[0055] S4. Based on the initial self-alignment error obtained in step S3 and the self-alignment error after alignment time t, the gyroscope zero-bias estimation result of the sub-inertial navigation system is obtained.

[0056] In this embodiment, in step S1, the analytical model of the relative installation error between the main inertial navigation system and the sub-inertial navigation system is constructed as follows:

[0057]

[0058] In the formula, Main inertial navigation in t m-1 to t m The angular increment vector at time t.

[0059] Main inertial navigation in t m-1 to t m The angular increment of the X-gyroscope at time t, Main inertial navigation in t m-1 to t m The angular increment of the Y-gyroscope at time t. Main inertial navigation in t m-1 to t m The angular increment of the Z-gyroscope at time t;

[0060] for antisymmetric matrix,

[0061]

[0062] The relative installation error vector between the master and sub-inertial navigation systems. Δβ 1,2 The pitch relative installation error of the master and slave inertial navigation systems, Δγ 1,2 The relative roll installation error of the master and sub-inertial navigation systems, Δα 1,2 The relative installation error of the heading of the main and sub-inertial navigation systems;

[0063] The relative attitude error vector caused by self-alignment error is determined by the following formula:

[0064]

[0065] In the formula, For t m The self-alignment error vector at time t.

[0066] δβ 1,2 (t m ) for t m The pitch relative attitude error at any given time, δγ 1，2 (t m ) for t m The relative roll attitude error at time δα 1，2 (tm ) for t m The relative attitude error of the heading at any given moment;

[0067] for antisymmetric matrix,

[0068]

[0069] For t m The relative attitude error vector at time t.

[0070] yes antisymmetric matrix,

[0071] In this embodiment, in step S2, the specific method for obtaining the relative installation error based on the analytical model of relative installation error constructed in step S1 includes the following steps:

[0072] S201. Collect the attitude vector ψ1 output by the main inertial navigation system, the attitude vector ψ2 output by the sub-inertial navigation system, and the angular increment vector output by the main inertial navigation system within a time interval T.

[0073] Where s≥3, ψ1=[β1,γ1,α1] T ψ2=[β2,γ2,α2] T β1 is the pitch angle output by the main inertial navigation system, γ1 is the roll angle output by the main inertial navigation system, α1 is the yaw angle output by the main inertial navigation system, β2 is the pitch angle output by the sub-inertial navigation system, γ2 is the roll angle output by the sub-inertial navigation system, and α2 is the yaw angle output by the sub-inertial navigation system.

[0074] S202. Based on the attitude vector ψ1 output by the main inertial navigation system and the attitude vector ψ2 output by the sub-inertial navigation system obtained in step S201, use the formula... Obtain the relative attitude error vector Right now

[0075]

[0076] In the formula, β1(t) m ) for t m The pitch angle output by the master inertial navigation system at time t, γ1(t) m ) for t m The roll angle output by the master inertial navigation system at time t, α1(t) m ) for t m The heading angle output by the master inertial navigation system at time t, β2(t) m ) for t m The pitch angle output by the sub-inertial navigation system at time t, γ2(t)m ) for t m The roll angle output by the sub-inertial navigation system at time t, α2(t) m ) for t m The heading angle output by the sub-inertial navigation system at any given moment;

[0077] S203. Based on step S1, construct an analytical model of the relative installation error between the main inertial navigation system and the sub-inertial navigation system, and use the attitude vectors output by the main inertial navigation system, the attitude vectors output by the sub-inertial navigation system, the angular increment vector output by the main inertial navigation system, and the relative attitude error vector obtained in step S201, and solve for the relative installation error vector.

[0078] Specifically, the attitude vectors output by the main inertial navigation system (INS), the attitude vectors output by the sub-INS, the angular increment vector output by the main INS, and the relative attitude error vectors obtained in step S201 within a time period T are substituted into the analytical model of relative installation error to obtain s sets of equations:

[0079]

[0080] An equation has three variables: Δβ 1,2 ,Δγ 1,2 and Δα 1,2 Since s≥3 is guaranteed in step S201, the solution is obtained by the least squares method. The relative installation error vector can then be obtained.

[0081] In this embodiment, the specific method for obtaining the initial self-alignment error and the self-alignment error after alignment time t using the relative installation error obtained in step S2 in step S3 includes the following steps:

[0082] S301. Collect the attitude vector ψ1(0) output at the initial moment after the main inertial navigation system is aligned and the attitude vector ψ2(0) output at the initial moment after the sub-inertial navigation system is aligned.

[0083] Where ψ1(0)=[β1(0),γ1(0),α1(0)] T ,ψ2(0)=[β2(0),γ2(0),α2(0)] T β1(0) is the pitch angle output at the initial moment of the main inertial navigation system's alignment completion, γ1(0) is the roll angle output at the initial moment of the main inertial navigation system's alignment completion, α1(0) is the heading angle output at the initial moment of the main inertial navigation system's alignment completion, β2(0) is the pitch angle output at the initial moment of the sub-inertial navigation system's alignment completion, γ2(0) is the roll angle output at the initial moment of the sub-inertial navigation system's alignment completion, and α2(0) is the heading angle output at the initial moment of the sub-inertial navigation system's alignment completion.

[0084] S302. Based on the attitude vector ψ1(0) output at the initial moment of the main inertial navigation system alignment completion and the attitude vector ψ2(0) output at the initial moment of the sub-inertial navigation system alignment completion obtained in step S301, using the formula... Obtain the self-alignment error vector at the initial moment after alignment completion, i.e.

[0085]

[0086] S303. Collect the attitude vector ψ1(t) output after time t after the main inertial navigation system alignment is completed, and the attitude vector ψ2(t) output after time t after the sub-inertial navigation system alignment is completed.

[0087] Among them, ψ1(t)=[β1(t),γ1(t),α1(t)] T ,ψ2(t)=[β2(t),γ2(t),α2(t)] T β1(t) is the pitch angle output after time t after the main inertial navigation system (INS) alignment is completed, γ1(t) is the roll angle output after time t after the main INS alignment is completed, α1(t) is the heading angle output after time t after the main INS alignment is completed, β2(t) is the pitch angle output after time t after the sub-INS alignment is completed, γ2(t) is the roll angle output after time t after the sub-INS alignment is completed, and α2(t) is the heading angle output after time t after the sub-INS alignment is completed.

[0088] S304. Based on the attitude vector ψ1(t) output after time t following the completion of the main inertial navigation system alignment and the attitude vector ψ2(t) output after time t following the completion of the sub-inertial navigation system alignment, obtained in step S303, the formula is used... Obtain the self-alignment error vector at the initial moment after alignment completion, i.e.

[0089]

[0090] In this embodiment, the specific method for obtaining the gyroscope zero-bias estimation result of the sub-inertial navigation system based on the initial self-alignment error obtained in step S3 and the self-alignment error after alignment time t includes the following steps in step S4:

[0091] S401. Acquire the attitude matrix output at the initial moment after the sub-inertial navigation alignment is completed.

[0092] S402, Using the formula Obtain the gyroscope zero-bias estimation result for the sub-inertial navigation system; where, The gyroscope zero-bias estimation result is given for the sub-inertial navigation system, where ΔT is the sensor sampling time, and t≥100ΔT.

[0093] To verify the correctness of this invention, a simulation experiment was conducted below.

[0094] In the simulation experiment, the performance parameters of the main inertial navigation system and the sub-inertial navigation system are listed in Table 1 and Table 2, respectively.

[0095] Table 1: Performance parameters of the main inertial navigation system in the simulation experiment

[0096] Main Inertial Navigation Zero bias stability noise Scale factor spinning top 0.01° / h 0.0015° / √h 10ppm accelerometer 20ug 2ug / √Hz 20ppm

[0097] Table 2: Performance parameters of the sub-inertial navigation system in the simulation experiment

[0098] Sub-inertial navigation Zero bias stability noise Scale factor spinning top 1° / h 0.1° / √h 300ppm accelerometer 300ug 30ug / √Hz 300ppm

[0099] Navigation data for the main and sub-inertial navigation systems were generated using the performance parameters corresponding to Tables 1 and 2. Then, the gyroscope zero bias of the sub-inertial navigation system was estimated using the method provided in this invention. A total of six sets of simulation experiments were conducted, with different simulated zero bias values ​​for each set. The experimental results are shown in Table 3. Table 3 lists the simulated zero bias values ​​for the main and sub-inertial navigation systems, the estimated zero bias values ​​for the sub-inertial navigation system, and the estimated residuals of the zero bias.

[0100] Table 3: Simulation results of zero-bias estimation for zero-bias compensation of the sub-inertial navigation system (unit: ° / h)

[0101]

[0102] As can be seen from the zero-bias estimation deviations in Table 3, when the zero-bias magnitude of the sub-inertial navigation system is 1° / h, the method provided by this invention can achieve a zero-bias estimation accuracy of better than 0.13° / h, effectively compensating for more than 80% of the zero-bias of the sub-inertial navigation system. Therefore, the effectiveness and correctness of the method provided by this invention are verified.

[0103] The parts of this invention not disclosed in detail are well-known technologies in the field.

[0104] Although the illustrative specific embodiments of the present invention have been described above to enable those skilled in the art to understand the invention, it should be understood that the invention is not limited to the scope of the specific embodiments. For those skilled in the art, various changes will be obvious as long as they are within the spirit and scope of the invention as defined and determined by the appended claims. All inventions utilizing the concept of the present invention are protected.

Claims

1. A method for estimating the zero bias of an airborne inertial navigation gyroscope combined with transfer alignment, characterized in that, Includes the following steps: S1. The analytical model of the relative installation error between the main inertial navigation system and the sub-inertial navigation system is as follows: ； In the formula, Main inertial navigation in t m-1 to t m The angular increment vector at time t. = , Main inertial navigation in t m-1 to t m The angular increment of the X-gyroscope at time t, Main inertial navigation in t m-1 to t m The angular increment of the Y-gyroscope at time t. Main inertial navigation in t m-1 to t m The angular increment of the Z-gyroscope at time t; for antisymmetric matrix, = ; The relative installation error vector between the master and sub-inertial navigation systems. Δβ 1,2 The pitch relative installation error of the master and slave inertial navigation systems, Δγ 1,2 The relative roll installation error of the master and sub-inertial navigation systems, Δα 1,2 The relative installation error of the heading of the main and sub-inertial navigation systems; The relative attitude error vector caused by self-alignment error is determined by the following formula: In the formula, For t m The self-alignment error vector at time t. = ,δβ 1,2 (t m ) for t m The pitch relative attitude error at any given time, δγ 1,2 (t m ) for t m The relative roll attitude error at time δα 1,2 (t m ) for t m The relative attitude error of the heading at any given moment; for antisymmetric matrix, = ; For t m The relative attitude error vector at time t. = ; yes antisymmetric matrix, = ； S2. Based on the analytical model of relative installation error constructed in step S1, obtain the relative installation error; S3. Using the relative installation error obtained in step S2, obtain the initial self-alignment error and the self-alignment error after alignment time t. S4. Based on the initial self-alignment error obtained in step S3 and the self-alignment error after alignment time t, the gyroscope zero-bias estimation result of the sub-inertial navigation system is obtained.

2. The method for estimating the zero bias of an airborne inertial navigation gyroscope combined with transfer alignment according to claim 1, characterized in that, In step S2, based on the analytical model of relative installation error constructed in step S1, the specific method for obtaining the relative installation error includes the following steps: S201. Collect s attitude vectors output by the main inertial navigation system within a time interval T. The attitude vector output by the sub-inertial navigation system and the angular increment vector output by the main inertial navigation system ; Where s≥3, , , The pitch angle output by the main inertial navigation system. The roll angle output by the main inertial navigation system. The heading angle output by the main inertial navigation system. The pitch angle output by the sub-inertial navigation system. The roll angle output by the sub-inertial navigation system. The heading angle output by the sub-inertial navigation system; S202. The attitude vector output by the main inertial navigation system obtained in step S201. attitude vector output by sub-inertial navigation Using the formula Obtain the relative attitude error vector ,Right now = = ; In the formula, for The pitch angle output by the main inertial navigation system at any given moment. for The roll angle output by the main inertial navigation system at any given time. for The heading angle output by the main inertial navigation system at any given time. for The pitch angle output by the sub-inertial navigation system at any given moment. for The roll angle output by the sub-inertial navigation system at any given moment. for The heading angle output by the sub-inertial navigation system at any given moment; S203. Based on step S1, construct an analytical model of the relative installation error between the main inertial navigation system and the sub-inertial navigation system, and use the attitude vectors output by the main inertial navigation system, the attitude vectors output by the sub-inertial navigation system, the angular increment vector output by the main inertial navigation system, and the relative attitude error vector obtained in step S201, and solve for the relative installation error vector. ; Specifically, the attitude vectors output by the main inertial navigation system (INS), the attitude vectors output by the sub-INS, the angular increment vector output by the main INS, and the relative attitude error vectors obtained in step S201 within a time period T are substituted into the analytical model of relative installation error to obtain s sets of equations: ； An equation has three variables: , and Since s≥3 is guaranteed in step S201, the solution is obtained by the least squares method. The relative installation error vector can then be obtained. .

3. The method for estimating the zero bias of an airborne inertial navigation gyroscope combined with transfer alignment according to claim 2, characterized in that, In step S3, the specific method for obtaining the initial self-alignment error and the self-alignment error after alignment time t using the relative installation error obtained in step S2 includes the following steps: S301. Acquire the attitude vector output at the initial moment after the main inertial navigation system alignment is completed. The attitude vector output at the initial moment after the sub-inertial navigation system alignment is completed. ; in, , , The pitch angle output at the initial moment after the main inertial navigation system's alignment ends. The roll angle output at the initial moment after the main inertial navigation system's alignment ends. The heading angle output at the initial moment after the main inertial navigation system ends alignment. The pitch angle output at the initial moment after the sub-inertial navigation system's alignment ends. The roll angle output at the initial moment after the sub-inertial navigation system's alignment is completed. The heading angle output at the initial moment after the sub-inertial navigation system's alignment is completed; S302. The attitude vector output at the initial moment after the alignment of the main inertial navigation system, obtained in step S301. The attitude vector output at the initial moment after the sub-inertial navigation system alignment is completed. Using the formula Obtain the self-alignment error vector at the initial moment after alignment completion, i.e. = ; S303. Acquire the attitude vector output t time after the main inertial navigation alignment is completed. The attitude vector output after time t following the completion of sub-inertial navigation alignment. ; in, , , The pitch angle output after time t following the completion of the main inertial navigation alignment. The roll angle output after time t following the completion of the main inertial navigation alignment. The heading angle output after time t following the completion of the main inertial navigation alignment. The pitch angle output after time t following the completion of the sub-inertial navigation alignment. The roll angle output after time t following the completion of sub-inertial navigation alignment. The heading angle output after time t following the completion of the sub-inertial navigation alignment; S304. The attitude vector output after time t following the completion of the main inertial navigation alignment, obtained in step S303. The attitude vector output after time t following the completion of sub-inertial navigation alignment. Using the formula Obtain the self-alignment error vector at the initial moment after alignment completion. Right now = .

4. The method for estimating the zero bias of an airborne inertial navigation gyroscope combined with transfer alignment according to claim 3, characterized in that: In step S4, based on the initial self-alignment error obtained in step S3 and the self-alignment error after alignment time t, the specific method for obtaining the gyroscope zero-bias estimation result of the sub-inertial navigation system includes the following steps: S401. Acquire the attitude matrix output at the initial moment after the sub-inertial navigation alignment is completed. ; S402, Using the formula Obtain the gyroscope zero-bias estimation result for the sub-inertial navigation system; where, The gyroscope bias estimation results for the sub-inertial navigation system. For sensor sampling time, .

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