A Method of Attitude Calculation for Inertial Navigation System
An inertial navigation system and attitude technology, applied in the field of inertial navigation system attitude calculation, can solve the problem of low accuracy of attitude calculation, achieve the effect of solving the problem of system attitude calculation and reducing the influence of attitude accuracy
Active Publication Date: 2022-04-12
TIANJIN NAVIGATION INSTR RES INST
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However, in practical applications, the acceleration of the carrier changes in real time, that is, the assumption of zero acceleration is not valid, resulting in low accuracy of attitude calculation
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[0045] The present invention will be described in further detail below in conjunction with the drawings and specific examples. The following examples are only descriptive, not restrictive, and cannot limit the protection scope of the present invention.
[0046] A kind of inertial navigation system attitude calculation method, see figure 1 , the solution method is based on the assumption of zero-mean acceleration, and a vehicle attitude update scheme based on the earth's fixed coordinate system is designed. First, select the carrier coordinate system b at time t=0 as the earth fixed coordinate system k 1 ,Right now Use the gyro output to update the attitude matrix, and obtain the carrier coordinate system b and the earth fixed coordinate system k within a given time interval 1 conversion relationship Then get the accelerometer in the earth fixed coordinate system k 1 Output According to the assumption of zero-mean acceleration, the earth fixed coordinate system k can be...
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The invention relates to a method for calculating the attitude of an inertial navigation system. Based on the assumption of zero-mean acceleration as a theoretical basis, a carrier attitude update scheme based on the earth-fixed coordinate system is designed; first, the carrier coordinate system b at time t=0 is selected as the earth-fixed joint coordinate system k 1 , that is, use the gyroscope output to update the attitude matrix, and obtain the carrier coordinate system b and the earth fixed coordinate system k in a given time interval 1 The conversion relationship of the accelerometer in the earth fixed coordinate system k 1 According to the zero-mean acceleration assumption, the output of the earth fixed coordinate system k can be obtained 1 system and the direction cosine matrix of the geographic coordinate system n combined with the carrier coordinate system b and the earth fixed coordinate system k 1 The transformation relationship can be obtained by the given time interval carrier coordinate system b and the direction cosine matrix of the geographic coordinate system n. Repeat the above process in the next given time interval to obtain the carrier attitude information. The invention effectively reduces the influence of the instantaneous acceleration of the carrier on the attitude accuracy, and better solves the problem of system attitude calculation under dynamic conditions.
Description
technical field [0001] The invention belongs to the technical field of strapdown inertial navigation systems, and in particular relates to an attitude calculation method of an inertial navigation system. Background technique [0002] The strapdown inertial navigation system includes a three-axis accelerometer and a three-axis gyroscope, and the horizontal attitude of the system can be obtained by using the output of the accelerometer and gyroscope. Traditional attitude calculation algorithms usually assume that the acceleration of the carrier motion is zero, that is, the accelerometer is only sensitive to the gravitational acceleration vector. The above assumptions are true when the carrier is static or moving at a constant speed, and accurate attitude information can be obtained through calculation. However, in practical applications, the acceleration of the carrier changes in real time, that is, the assumption of zero acceleration is not valid, resulting in low accuracy o...
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IPC IPC(8): G01C21/18G01C21/16G01C25/00
CPCG01C21/18G01C21/16G01C25/005
Inventor 杨松普王琳
Owner TIANJIN NAVIGATION INSTR RES INST