High-precision autonomous navigation system based on miniaturized indexing mechanism

Abstract

The invention discloses a high-precision autonomous navigation system based on a miniaturized indexing mechanism. An inertial measurement assembly is used for measuring angular velocity and acceleration information of an aircraft; in the double-shaft indexing mechanism, an inner frame shaft system adopts a 73LST sine wave to drive a permanent magnet torque motor so as to drive the inner frame shaft system to move and position, and a J75XFS absolute type rotary transformer is adopted as an angle measurement and angular position feedback element of the shaft system; the outer frame shaft system adopts a J85LST sine wave driving permanent magnet torque motor and a J75XFS absolute type rotary transformer to jointly complete the displacement and positioning of the outer frame shaft system; the locking device is arranged in the direction perpendicular to the rotation axis of the inner ring shaft system and the outer ring shaft system. The electrical control system comprises a spin lock circuit, two ends of the inner and outer frame shaft systems are respectively provided with a torque motor and a multi-stage resolver goniometer, and the torque motor and the multi-stage resolver goniometer and the spin lock circuit of the electrical control system form a closed-loop control loop; and the navigation software module is used for realizing high-precision autonomous navigation of the aircraft.

Description

Technical Field

[0001] The invention relates to a high-precision autonomous navigation system based on a miniaturized transfer mechanism, belonging to the technical field of inertial navigation. Background Art

[0002] The existing traditional strapdown positioning and orientation inertial navigation system is based on the inertial navigation system and the satellite navigation system. It can measure the acceleration and angular velocity of the three axes of the spatial coordinate system for positioning and orientation, and achieve high-precision navigation by combining satellite navigation positioning. Some aircraft or spacecraft achieve autonomous navigation by adding high-speed rotating components, star sensors and high-precision digital maps, but each method has its drawbacks. Not only will the cost increase, but there are also limitations in use, which makes it difficult to meet the high-precision autonomous navigation needs of aircraft or spacecraft. In addition, the current inertial group adopts a self-calibration scheme, which requires setting a reasonable calibration path and using the existing calibration model to compensate for the gyroscope and accelerometer output, resulting in a long calibration time and cannot meet the requirements of fast calibration. Summary of the invention

[0003] The technical problem to be solved by the present invention is to overcome the deficiencies of the prior art and provide a high-precision autonomous navigation system based on a miniaturized transfer mechanism, which does not rely on satellite navigation and can achieve rapid calibration of an inertial group.

[0004] The purpose of the present invention is achieved through the following technical solutions:

[0005] A high-precision autonomous navigation system based on a miniaturized indexing mechanism realizes online suppression of the equivalent zero position of the inertial group through a rotation modulation technology based on a dual-axis indexing device, thereby achieving high-precision autonomous navigation of small aircraft.

[0006] The details are as follows:

[0007] (1) Miniaturized dual-axis indexing mechanism

[0008] The high-precision autonomous navigation system consists of an inertial measurement unit, a dual-axis transfer mechanism, and a navigation software module. The dual-axis transfer mechanism is mainly composed of an inner frame axis system, an outer frame axis system, a locking device, and an electrical control system. In order to achieve product miniaturization, the inner frame shaft system adopts 73LST sinusoidal wave driven permanent magnet torque motor to drive the inner frame shaft system for movement and positioning, and adopts J75XFS absolute rotary transformer as the angle measurement and angular position feedback element of the shaft system; the outer frame shaft system adopts J85LST sinusoidal wave driven permanent magnet torque motor and J75XFS absolute rotary transformer to jointly complete the movement and positioning of the outer frame shaft system; the locking device is arranged in the vertical direction of the rotation axis of the inner and outer ring shaft systems, and a gear locking scheme is adopted. When the inner and outer rings are both at 0°, the IMU is locked and unlocked by controlling the lifting and lowering of the movable gear disc. Compared with the traditional scheme, the volume of this scheme is reduced by more than 40%; the electrical control system includes a rotary lock circuit and related circuits. Torque motors and multi-stage rotary angle meters are installed at both ends of the inner and outer frame shaft systems, respectively, and a closed-loop control circuit is formed with the rotary lock circuit to jointly control the frame shaft system to drive the rotation and positioning of the inner platform.

[0009] The inertial measurement unit is mainly composed of a three-axis fiber optic gyroscope, a three-axis quartz flexible accelerometer and an IF conversion circuit, and is used to measure the angular velocity and acceleration information of the aircraft.

[0010] The navigation software module realizes online suppression of inertial group errors, high-precision attitude alignment, and online rapid estimation and compensation of equivalent zero position based on rotational modulation technology, thereby achieving high-precision autonomous navigation of the aircraft.

[0011] (2) A high-precision inertial group equivalent zero position fast estimation method based on a three-inertial group

[0012] In the system framework of the present invention, the navigation software module can realize online suppression of inertial group errors, high-precision attitude alignment, and online rapid estimation and compensation of equivalent zero position based on rotation modulation technology, thereby realizing high-precision autonomous navigation of the aircraft, as follows:

[0013] a) Establish the inertial group error model, the indexing mechanism error model and the rotation modulation scheme, and realize the online suppression of the inertial group error based on the rotation modulation inertial group error suppression technology, specifically:

[0014] Assume that the carrier coordinate system coincides with the IMU rotation coordinate system at the initial moment, and control the IMU to rotate continuously around the Y axis at ω. The modulation form of the gyro constant drift and accelerometer constant zero bias in the navigation coordinate system is:

[0015]

[0016]

[0017] in, is the transformation matrix between the carrier coordinate system and the navigation coordinate system; It is the transformation matrix between the IMU rotation coordinate system and the carrier coordinate system; is the gyro constant drift in three axes; is the constant zero bias of the accelerometer in the three axes.

[0018] Therefore, when the IMU rotates, taking the gyro equivalent eastward constant drift as an example, it can be expressed as

[0019]

[0020] In the formula, When the IMU rotates a full circle, the equation sin(ωt+φ)-sinφ=0 is satisfied, which can completely eliminate the influence of drift terms such as the equivalent gyro constant drift, thereby achieving online suppression of the aircraft inertial group error.

[0021] b) High-precision attitude alignment is achieved through rotation modulation based on high-precision attitude initial alignment technology to reduce the initial attitude error of the aircraft. Specifically:

[0022] The initial alignment error equation for a single position is established to obtain the equivalent zero bias terms of the gyro and accelerometer. By analyzing the error terms in the initial alignment error equation, it can be concluded that as long as the influence of the gyro zero bias and the accelerometer zero bias can be reduced, the initial alignment accuracy can be improved. The rotation modulation technology can be used to compensate for the initial attitude error. By using its sinusoidal characteristics, the error terms of the gyro and accelerometer can be offset after a full rotation of the IMU. After the above analysis, the three-inertial group can achieve high-precision attitude alignment through rotation modulation, reduce the initial attitude error of the aircraft, and improve the navigation accuracy of the aircraft.

[0023] c) Through rotation modulation based on the inertial group equivalent zero position online fast estimation and compensation technology to achieve online estimation and compensation of the gyro and accelerometer equivalent zero position, reducing the cumulative error

[0024] Aiming at the problem that the calibration time of inertial group error is too long, the present invention proposes an online rapid estimation and compensation technology of equivalent zero position, which is used for rapid estimation, calibration and compensation of error. It is necessary to establish an inertial group error model under a rotating coordinate system, that is, to derive the attitude error equation and the velocity error equation of the self-calibrated inertial navigation system according to the error equation of the strapdown inertial navigation system, and use Kalman filtering to estimate the equivalent zero position of the inertial group online. Aiming at the problem that when the accuracy of the inertial group is high, the equivalent zero position of the gyro is small and difficult to estimate or the estimation accuracy is poor, the fourth-order Runge Kutta method is used to solve the formula, and the attitude error of the inertial navigation stage is stimulated to achieve accurate estimation of the equivalent zero position of the gyro.

[0025] Utilizing the characteristics of the sinusoidal function, the inertial navigation system error suppression technology based on rotation modulation can completely eliminate the influence of the equivalent zero position, realize online rapid estimation and compensation of the inertial navigation system equivalent zero position, and greatly improve navigation accuracy.

[0026] Before the aircraft takes off, the miniaturized indexing mechanism can calibrate its constant zero bias and scale coefficient error based on the above technology. The speed can be increased to within 40s, and the requirements for high-precision inertial group error estimation can be met. Rotational modulation technology is used for compensation, and the zero bias stability can reach 1 / 60 of the original parameters.

[0027] Compared with the prior art, the present invention has the following beneficial effects:

[0028] (1) The high-precision autonomous navigation system based on a miniaturized transfer mechanism improves navigation accuracy through rotation modulation technology. When the GNSS is interfered with (or fails), the system can achieve high-precision autonomous navigation to achieve subsequent precise guidance.

[0029] (2) The design of a miniaturized dual-axis indexing mechanism is based on a three-inertial group module to achieve high-precision autonomous navigation of the system, and the volume can be reduced by 40%, meeting the miniaturization requirements of the aircraft.

[0030] (3) The high-precision inertial group equivalent zero position fast estimation technology based on the three-self inertial group can increase the estimation speed of the inertial group equivalent zero position to less than 40 seconds, meeting the needs of high-precision inertial group error estimation.

[0031] (4) The three-auto inertial navigation module can improve the inertial navigation accuracy by suppressing the inertial navigation error, improving the initial alignment accuracy of the navigation, estimating the equivalent zero bias of the inertial navigation and compensating for it. The inertial navigation accuracy can be improved by more than 80%. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] Figure 1 This is a structural diagram of the high-precision autonomous navigation system based on a miniaturized indexing mechanism of the present invention;

[0033] Figure 2 It is the gyro bias and standard bias curve of the high-precision autonomous navigation system based on the miniaturized indexing mechanism of the present invention;

[0034] Figure 3 The navigation angle, pitch angle and roll angle error curves of the high-precision autonomous navigation system based on the miniaturized indexing mechanism of the present invention;

[0035] Figure 4 The rotation modulation alignment position error curve of the high-precision autonomous navigation system based on the miniaturized indexing mechanism of the present invention;

[0036] Figure 5 The equivalent zero bias estimation value curve of the gyro and table of the high-precision autonomous navigation system based on the miniaturized indexing mechanism of the present invention;

[0037] Figure 6 This is the zero bias correction combined position error curve of the high-precision autonomous navigation system based on the miniaturized indexing mechanism of the present invention. DETAILED DESCRIPTION

[0038] To make the objectives, technical solutions and advantages of the present invention more clear, the embodiments of the present invention will be further described in detail below with reference to the accompanying drawings.

[0039] A high-precision autonomous navigation system based on a miniaturized indexing mechanism, such as Figure 1 As shown, it consists of a dual-axis transfer mechanism, an inertial measurement unit, and a navigation software module.

[0040] The dual-axis indexing mechanism is mainly composed of inner and outer frame shaft systems and an electrical control system. In order to achieve product miniaturization, the inner frame shaft system adopts a 73LST sinusoidal wave driven permanent magnet torque motor to drive the inner frame shaft system for movement and positioning, and adopts a J75XFS absolute rotary transformer as the angle measurement and angular position feedback element of the shaft system; the outer frame shaft system adopts a J85LST sinusoidal wave driven permanent magnet torque motor and a J75XFS absolute rotary transformer to jointly complete the movement and positioning of the outer frame shaft system; the locking device is arranged in the vertical direction of the rotation axis of the inner and outer ring shaft systems, and a toothed disc locking scheme is adopted. When the inner and outer rings are both at 0°, the IMU is locked and unlocked by controlling the lifting and lowering of the movable toothed disc. Compared with the traditional scheme, the volume of this scheme is reduced by more than 40%; the electrical control system includes a rotary lock circuit and related circuits, and torque motors and multi-stage rotary angle measuring devices are respectively installed at both ends of the inner and outer frame shaft systems, and a closed-loop control circuit is formed with the rotary lock circuit to jointly control the frame shaft system to drive the rotation and positioning of the inner platform.

[0041] The inertial measurement unit is mainly composed of a three-axis fiber optic gyroscope, a three-axis quartz flexible accelerometer and an IF conversion circuit, and is used to measure the angular velocity and acceleration information of the aircraft.

[0042] The navigation software is based on the rotation modulation technology to achieve online suppression of inertial group errors, high-precision attitude alignment, and online rapid estimation and compensation of equivalent zero position, thereby achieving high-precision autonomous navigation of the aircraft, as follows:

[0043] (a) Establish the inertial group error model, the indexing mechanism error model and the rotation modulation scheme, and realize the online suppression of the inertial group error based on the rotation modulation inertial group error suppression technology.

[0044] Establish a model construction plan, assume that the carrier coordinate system coincides with the IMU rotation coordinate system at the initial moment, control the IMU to rotate continuously around the Y axis at ω, and the relationship between the carrier system and the IMU rotation coordinate system can be obtained as follows:

[0045]

[0046] Assume that the gyro constant drift of the three axes is The constant bias of the accelerometer is In order to analyze the mechanism of rotation modulation on the constant error of inertial devices intuitively, assuming that the navigation coordinate system coincides with the carrier coordinate system, the modulation form of the gyro constant drift and accelerometer constant zero bias in the navigation coordinate system at time t is:

[0047]

[0048] In the formula, is the transformation matrix between the carrier coordinate system and the navigation coordinate system; It is the transformation matrix between the IMU rotation coordinate system and the carrier coordinate system.

[0049] Therefore, when the IMU rotates, taking the gyro equivalent eastward constant drift as an example, it can be expressed as

[0050]

[0051] In the formula, It can be seen that when the north component of the Earth's rotation angular velocity ω N When determining, the greater the IMU rotation angular velocity, the smaller the horizontal gyro constant drift error, the longer the rotation alignment time, and the higher the azimuth alignment accuracy.

[0052] When the IMU rotates a full circle, the equation sin(ωt+φ)-sinφ=0 is satisfied, so the influence of the equivalent east gyro constant drift can be completely eliminated. Similarly, the equivalent north gyro zero bias Equivalent east accelerometer bias and equivalent north accelerometer bias After the IMU rotates a full circle, it can also be offset. In addition, the continuous rotation of the IMU around the Y axis can change the system matrix in the SINS error model, thereby improving the observability of the system and the observability of the state variables, thereby improving the alignment accuracy and speed.

[0053] (b) High-precision attitude alignment is achieved through rotation modulation based on high-precision attitude initial alignment technology to reduce the initial attitude error of the aircraft

[0054] In the initial alignment of a single position, the estimated error of the initial alignment misalignment angle satisfies the following relationship:

[0055]

[0056] Where δφ is the estimated error of the misalignment angle; is the equivalent east accelerometer bias; is the equivalent north accelerometer bias; is the equivalent east gyro zero bias; ω N is the north component of the earth's rotational angular velocity; g is the magnitude of gravitational acceleration; L is the geographic latitude. In general, the measurement error of the gyroscope relative to the earth's rotation is significantly greater than the measurement error of the accelerometer relative to the earth's gravity, that is,

[0057]

[0058] Therefore, the equivalent east gyro bias of the gyro It is the main factor affecting the azimuth misalignment angle (yaw) estimation error, that is, the yaw angle alignment accuracy mainly depends on The pitch and roll alignment accuracy depends mainly on the accelerometer and

[0059] The rotation modulation technology is used to compensate the initial attitude error. When the IMU rotates a full circle, the equation sin(ωt+φ)-sinφ=0 is still satisfied, completely eliminating the influence of the equivalent east gyro zero bias, that is, and After the IMU rotates a full circle, it can be offset.

[0060] According to the above method description, it can be concluded that as long as the and The influence of can improve the initial alignment accuracy. After the above analysis, the three-inertial group can achieve high-precision attitude alignment through rotation modulation and reduce the initial attitude error of the aircraft.

[0061] (c) Through rotation modulation based on the online fast estimation and compensation technology of the equivalent zero position of the inertial group, the online estimation and compensation of the equivalent zero position of the gyroscope and accelerometer can be realized to reduce the cumulative error

[0062] Gyroscopes are inertial devices that are sensitive to the angular velocity of moving carriers. Gyroscope drift, scale factor error, and installation error mainly affect the carrier's attitude solution. Accelerometers are inertial devices that are sensitive to the acceleration of moving carriers. Acceleration zero bias, scale factor error, and installation error affect the carrier's velocity solution, and coupled with the angular velocity, affect the navigation system's position solution.

[0063] Gyro drift mainly includes slow drift and random quantity, among which the slow drift component will change slowly during operation. The drift of gyro startup depends on the random factors at the startup time. When the gyro is started, the drift can be defined as a fixed constant, so it is regarded as gyro constant drift. Gyro random drift changes randomly during operation, and the gyro scale coefficient error can also be described as a random constant, which is regarded as a fixed constant after a single startup. The installation error is generally defined as a constant after the gyro leaves the factory.

[0064] Similarly, accelerometer errors mainly include constant zero bias, scale coefficient error, installation error and random zero bias. The nature of their error values ​​is similar to gyroscope drift, and both need to be calibrated and compensated before launch.

[0065] There are two traditional methods for estimating the inertial group error based on the three-self-inertial group. The first is the self-calibration scheme. This scheme requires setting a reasonable calibration path and then using the existing calibration model to compensate for the gyro and accelerometer outputs. This scheme takes a long time to calibrate and cannot meet the requirements of rapid calibration and launch of aircraft or spacecraft. The second is the inertial group equivalent zero position estimation method based on rotation modulation technology. This method has insufficient estimation accuracy for the high-precision inertial group equivalent zero position. In response to the above problems, this scheme proposes a high-precision inertial group equivalent zero position fast estimation technology based on the three-self-inertial group. Compared with the traditional scheme, the estimation speed of the inertial group equivalent zero position is increased from 50 minutes to within 40 seconds. At the same time, it can be applied to the field of high-precision inertial group error estimation. The details are as follows:

[0066] The outputs of the gyro and accelerometer of the rotating inertial navigation system are and f p , and its error is expressed as and δf p During navigation solution, the data measured in the p system is converted to the b system for solution after rotation decomposition, and the measurement adds one coordinate transformation. According to the error equation of the strapdown inertial navigation system, the attitude error equation and velocity error equation of the self-calibration inertial navigation system can be derived, as follows:

[0067]

[0068] In the above formula, n represents the navigation coordinate system, b represents the carrier coordinate system, p represents the rotating coordinate system, i represents the inertial coordinate system, and e represents the earth coordinate system. is the misalignment angle of the mathematical platform, v and δv are the carrier velocity and velocity error, ω and δω are the carrier angular velocity and angular velocity error, f and δf are the carrier specific force and its error, δg is the gravity deviation, is the transformation matrix from the carrier coordinate system to the navigation coordinate system, is the transformation matrix from the rotation coordinate system to the carrier coordinate system, is the error between the rotating system and the carrier system. is the projection of the angular velocity of the navigation coordinate system relative to the inertial coordinate system in the navigation coordinate system, f n is the specific force in the navigation coordinate system, δf p is the specific force error in the navigation coordinate system, is the projection of the angular velocity of the earth coordinate system relative to the inertial coordinate system in the navigation coordinate system, is the projection of the angular velocity of the navigation coordinate system relative to the earth coordinate system in the navigation coordinate system, Expressed as error, Expressed as error.

[0069] When the accuracy of the inertial navigation system is high, δω and δf are small quantities and are difficult to estimate or have poor estimation accuracy. To address this problem, this solution proposes a high-precision inertial navigation system equivalent zero position fast estimation technology based on a three-self inertial navigation system. By stimulating the attitude and velocity errors in the inertial navigation stage, accurate estimation of δω and δf is achieved. The details are as follows:

[0070] In strapdown navigation, the attitude of the carrier system to the navigation system is required, and the following quaternion motion equations must be solved:

[0071]

[0072] in, and It is the projection of the angular velocity of the three-axis aircraft or spacecraft body coordinate system relative to the navigation coordinate system in the aircraft or spacecraft body coordinate system, and Q is a quaternion.

[0073] This system uses the fourth-order Longo Kutta method to solve the equation, as follows:

[0074]

[0075] In the formula, T is the step size of attitude solution. In order to stimulate δω, T is magnified by 100 times, and δω is correspondingly magnified by 100 times. After magnification, δω can be estimated and compensated with high precision. Similarly, this scheme can be extended to other magnification factors.

[0076] Based on the high-precision inertial group equivalent zero position fast estimation technology of the above three-inertial group, before the launch of the aircraft or spacecraft, the three-inertial group can calibrate its constant zero bias and scale coefficient error, and compensate it through rotation modulation technology. It has been verified that in the case of uninterrupted power, the navigation system error is only random zero bias, that is, random walk. The test results show that the zero bias stability after 1s averaging can reach 1 / 60 of the original parameter, greatly improving the navigation accuracy.

[0077] In order to further illustrate the above method and fully demonstrate the reliability and accuracy of the method, a digital simulation experiment was carried out.

[0078] (1) Description of simulation conditions and parameter settings

[0079] Simulation conditions: gyro bias 0.5° / h, gyro bias stability 0.3° / h, gyro bias random walk 0.05° / h, table bias 0.3mg, bias stability 0.3mg, table bias random walk 0.03mg;

[0080] (2) Simulation test results and analysis

[0081] (a) Inertial group error suppression technology based on rotation modulation

[0082] Figure 2 The simulation test results of the present invention to verify the effectiveness of the rotation modulation scheme are given. It can be seen that the horizontal gyro and the constant drift of the adder are modulated into a sine wave with a period average of 0, and the influence of the equivalent gyro constant drift can be completely offset by rotation (IMU rotation of a full circle). As shown in Table 1, the three-self inertial group rotation modulation can greatly reduce the gyro equivalent zero bias and the equivalent zero bias of the adder, thereby realizing the online suppression of the inertial group error.

[0083] Table 1 Comparison of equivalent zero bias of gyro and adding table before and after rotation modulation

[0084] Normal value After rotation modulation Remark Gyro equivalent zero bias (° / h) 1 0.0029 Add table equivalent zero bias (mg) 0.3 0.006

[0085] (b) High-precision posture initial alignment technology

[0086] Figures 3-4 The simulation test results of the present invention to verify the influence of the rotation modulation scheme on the initial posture error are given. Figure 3 As shown in Table 2, the comparison of navigation angle, pitch angle, and roll angle error values ​​is given; Figure 4 As shown in the figure, in the normal strapdown alignment mode, the position error is 4079m, and in the rotational modulation alignment mode, the position error is 517.4m, which is 12.7% of the error in the normal strapdown alignment mode. Comprehensive analysis shows that the rotational modulation technology can achieve high-precision attitude alignment and reduce the initial attitude error of the aircraft.

[0087] Table 2 Comparison of alignment accuracy between traditional strapdown scheme and rotation modulation scheme

[0088] Traditional strapdown solution Rotary modulation scheme Remark Pitch angle error (°) 0.017 <![CDATA[3.41×10 -4 ]]> Roll angle error (°) 0.017 <![CDATA[2.04×10 -4 ]]> Heading angle error (°) 8.30 (external input, accuracy 0.2) 0.01

[0089] (c) Online fast estimation and compensation technology of equivalent zero position of inertial group

[0090] Figures 5-6 The simulation test results of the present invention to verify the influence of the rotation modulation scheme on the equivalent zero bias are given. Figure 5 As shown in , the efficiency of the gyro bias estimation is about 80%, and the efficiency of the table bias estimation is about 70%, that is, the equivalent zero position of the gyro and accelerometer is well estimated; Figure 6As shown in the figure, the navigation position error of the rotation modulation alignment plus zero bias correction method is 214.1m, which is 41.4% of the rotation modulation alignment and 5.2% of the traditional strapdown inertial navigation. Comprehensive analysis shows that the rotation modulation technology can realize the online estimation and compensation of the equivalent zero position of the gyroscope and accelerometer, and reduce the cumulative error.

[0091] The results show that the high-precision autonomous navigation system based on the miniaturized transfer mechanism of the present invention utilizes the rotation modulation of the three-inertial group to suppress the inertial group error, estimates the equivalent zero bias and compensates it. Compared with the traditional strapdown navigation method, it greatly reduces the error and can effectively improve the navigation accuracy of aircraft or spacecraft.

[0092] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

[0093] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A high-precision autonomous navigation system based on a miniaturized indexing mechanism, characterized in that: Includes inertial measurement unit, dual-axis indexing mechanism, and navigation software module; The inertial measurement unit is used to measure the angular velocity and acceleration information of the aircraft; The dual-axis indexing mechanism consists of an inner frame shaft system, an outer frame shaft system, a locking device and an electrical control system. The inner frame shaft system adopts a 73LST sinusoidal wave driven permanent magnet torque motor to drive the inner frame shaft system for movement and positioning, and adopts a J75XFS absolute rotary transformer as the shaft system angle measurement and angular position feedback element; the outer frame shaft system adopts a J85LST sinusoidal wave driven permanent magnet torque motor and a J75XFS absolute rotary transformer to complete the movement and positioning of the outer frame shaft system; the locking device is set in the vertical direction of the rotation axis of the inner and outer ring shaft systems, and a gear disc locking scheme is adopted; the electrical control system includes a rotary lock circuit, and torque motors and multi-stage rotary angle measuring devices are installed at both ends of the inner and outer frame shaft systems respectively, and form a closed-loop control circuit with the rotary lock circuit; The navigation software module realizes online suppression of inertial group errors, high-precision attitude alignment, and online rapid estimation and compensation of equivalent zero position based on rotational modulation, thereby achieving high-precision autonomous navigation of the aircraft.

2. The high-precision autonomous navigation system according to claim 1, characterized in that: The locking device adopts a gear locking solution. When the inner and outer rings are both at 0°, the IMU is locked and unlocked by controlling the lifting and lowering of the movable gear.

3. The high-precision autonomous navigation system according to claim 1, characterized in that: In the navigation software module, it is assumed that the carrier coordinate system coincides with the IMU rotation coordinate system at the initial moment, and the IMU is controlled to rotate continuously around the Y axis at ω. The relationship between the carrier system and the IMU rotation coordinate system is: Assume that the gyro constant drift of the three axes is The constant bias of the accelerometer is Assuming that the navigation coordinate system coincides with the carrier coordinate system, the modulation form of the gyro constant drift and accelerometer constant zero bias in the navigation coordinate system at time t is: In the formula, is the transformation matrix between the carrier coordinate system and the navigation coordinate system; It is the transformation matrix between the IMU rotation coordinate system and the carrier coordinate system.

4. The high-precision autonomous navigation system according to claim 3, characterized in that: When the IMU rotates, take the gyro equivalent eastward constant drift as an example: In the formula, When the north component of the Earth's rotational angular velocity ω N When determining, the greater the IMU rotation angular velocity, the smaller the horizontal gyro constant drift error, the longer the rotation alignment time, and the higher the azimuth alignment accuracy.

5. The high-precision autonomous navigation system according to claim 4, characterized in that: In the navigation software module, when the IMU rotates a full circle, the equation sin(ωt+φ)-sinφ=0 is satisfied, thus completely eliminating the influence of the equivalent eastward gyro constant drift.

6. The high-precision autonomous navigation system according to claim 5, characterized in that: Equivalent north gyro bias Equivalent east accelerometer bias and equivalent north accelerometer bias It can also be offset after the IMU rotates a full circle.

7. The high-precision autonomous navigation system according to claim 4, characterized in that: In the navigation software module, high-precision attitude alignment is achieved through rotation modulation based on high-precision attitude initial alignment technology to reduce the initial attitude error of the aircraft.

8. The high-precision autonomous navigation system according to claim 7, characterized in that: In the initial alignment of a single position, the estimated error of the initial alignment misalignment angle satisfies the following relationship: Where δφ is the estimated error of the misalignment angle; is the equivalent east accelerometer bias; is the equivalent north accelerometer bias; is the equivalent east gyro bias; ω N is the north component of the earth's rotational angular velocity; g is the magnitude of gravitational acceleration; L is the geographic latitude; the measurement error of the gyroscope relative to the earth's rotation is significantly greater than the measurement error of the accelerometer relative to the earth's gravity, that is, Therefore, the equivalent east gyro bias of the gyro It is the main factor affecting the azimuth misalignment angle (yaw) estimation error, that is, the yaw angle alignment accuracy mainly depends on The pitch and roll alignment accuracy depends mainly on the accelerometer and The rotation modulation technology is used to compensate the initial attitude error. When the IMU rotates a full circle, the equation sin(ωt+φ)-sinφ=0 is still satisfied, completely eliminating the influence of the equivalent east gyro zero bias, that is, and After the IMU rotates a full circle, it can be offset.

9. The high-precision autonomous navigation system according to claim 8, characterized in that: The three-inertia group reduces the and and improve the initial alignment accuracy.

10. The high-precision autonomous navigation system according to claim 4, characterized in that: Based on the online fast estimation and compensation technology of the equivalent zero position of the inertial group, the online estimation and compensation of the equivalent zero position of the gyroscope and accelerometer are realized to reduce the cumulative error.

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