A Method of Inertial Navigation System Attitude Estimation

An attitude estimation and inertial navigation system technology, applied in the field of inertial navigation system attitude estimation, can solve the problems of interference measurement, inaccurate horizontal attitude estimation, large errors, etc., and achieve the effect of reducing interference, high engineering application value, and optimal estimation

Active Publication Date: 2021-12-31
TIANJIN NAVIGATION INSTR RES INST
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Problems solved by technology

However, when the carrier has a motion acceleration, the accelerometer output includes the local gravity acceleration and motion acceleration. At this time, if the gravity vector is used as a reference to calculate the horizontal attitude, it will lead to a large error
That is, the acceleration of the carrier motion interferes with the measurement of the gravity vector by the accelerometer, making the horizontal attitude estimation inaccurate

Method used

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  • A Method of Inertial Navigation System Attitude Estimation
  • A Method of Inertial Navigation System Attitude Estimation
  • A Method of Inertial Navigation System Attitude Estimation

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Embodiment Construction

[0051] Embodiments of the present invention are described in further detail below in conjunction with the accompanying drawings:

[0052] An inertial navigation system attitude estimation method, such as figure 1 shown, including the following steps:

[0053] Step 1, taking the attitude quaternion as the state quantity, and establishing the Kalman filter state equation on the basis of the quaternion attitude update differential equation;

[0054] The concrete steps of described step 1 include:

[0055] (1) According to the known accelerometer output is y a =[0a x a y a z ]T , the gyroscope output is y g =[0 ω x ω y ω z ] T , the gravity vector is G=[0 0 0 -g] T , to establish the Kalman filter state equation, the known basic form of the state equation is:

[0056] x k+1 = Φ k x k +W k

[0057] Among them, state quantity X=[q 0 q 1 q 2 q 3 ε x ε y ε z ] T , q is the carrier attitude quaternion, ε is the gyro constant drift; state transition mat...

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Abstract

The present invention relates to a kind of attitude estimation method of inertial navigation system, and its technical feature is: comprise the following steps: Step 1, take attitude quaternion as state quantity, establish Kalman filter state equation on the basis of quaternion attitude updating differential equation; Step 2, set up the quaternion pseudo Kalman wave measurement equation with the accelerometer measurement value as input; Step 3, the Kalman filter state equation established according to step 1 and the quaternion pseudo Kalman wave measurement equation established in step 2 Carry out Kalman filter measurement update, calculate the measurement update residual value, and estimate the measurement noise variance in real time according to the accelerometer measurement update residual value, dynamically adjust the size of the measurement noise variance matrix through the size of the motion acceleration, Then the attitude measurement accuracy of the system under dynamic conditions can be guaranteed. The invention effectively improves the estimation accuracy of the horizontal posture under dynamic conditions.

Description

technical field [0001] The invention belongs to the technical field of IMU attitude measurement, in particular to an inertial navigation system attitude estimation method. Background technique [0002] An IMU-based attitude measurement system that utilizes accelerometer and gyroscope outputs for optimal estimation of horizontal attitude. When the carrier has no motion acceleration, the accelerometer can accurately sense the local gravity vector and fuse with the output of the gyroscope to obtain a higher-precision horizontal attitude. However, when the carrier has motion acceleration, the accelerometer output includes the local gravity acceleration and motion acceleration. At this time, if the gravity vector is used as a reference to calculate the horizontal attitude, it will lead to a large error. That is, the acceleration of the carrier motion interferes with the measurement of the gravity vector by the accelerometer, making the horizontal attitude estimation inaccurate. ...

Claims

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Application Information

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Patent Type & Authority Patents(China)
IPC IPC(8): G01C9/00G01C21/18
CPCG01C9/00G01C21/18
Inventor 杨松普王琳宋高玲赵汪洋
Owner TIANJIN NAVIGATION INSTR RES INST
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