Abstract
A method of accurate integrated navigation for high-altitude aerocraft by medium precision strapdown inertial navigation system (SINS), star sensor, and global navigation satellite system (GNSS) is researched in this paper. The system error sources of SINS and star sensor are analyzed and modeled, and then system errors of SINS and star sensor are chosen as system states of integrated navigation. Considering that the output of star sensor is attitude quaternion, it can be regarded as an attitude matrix, then the equivalent attitude matrix is constructed by using the output of SINS, and the calculating equation of the equivalent attitude matrix is designed. Thus, one of the measurements of integrated navigation can be constructed by using the equivalent attitude matrix and the attitude matrix output of star sensor. According to the constraint conditions of the attitude matrix, the diagonal elements are selected as one of the measurements of integrated navigation, and the corresponding measurement equation is derived. At the same time, the velocity output and position output difference between SINS and GNSS is selected as the other measurement, and the corresponding measurement equation is also derived. On this basis, the Kalman filter is used to design an integrated navigation filtering algorithm. Simulation results show that although the medium precision SINS is used, the heading accuracy of this integrated navigation method is better than ±1.5′, the pitch and roll accuracy are better than ±0.9’, the velocity accuracy is better than ±0.05 m/s, and the position accuracy is better than ±3.8 m. Therefore, the integrated navigation effect is very significant.
Highlights
For the flight control of high-altitude aerocraft, the navigation ability with high precision and high reliability is very important; especially in the long endurance environment, it is not easy to be affected by external natural environment and human interference
Due to the fact that satellite navigation accuracy is much higher than other single navigation technologies, the errors of global navigation satellite system (GNSS) are directly considered as a white noise process, which are not chosen as system states of integrated navigation anymore
In order to construct the measurement of integrated navigation, it is necessary to calculate the equivalent attitude matrix according to the outputs of strapdown inertial navigation system (SINS) and other pieces of information, which is denoted as mC bia.trixisC bieqoufivstaalernstenatstoitru. de matrix can match the attitude
Summary
For the flight control of high-altitude aerocraft, the navigation ability with high precision and high reliability is very important; especially in the long endurance environment, it is not easy to be affected by external natural environment and human interference. Because of its high attitude determination accuracy, star sensor can be used to assist INS/satellite-integrated navigation system to improve its heading accuracy and gyro drift estimation effect and can significantly improve the antijamming ability of the whole integrated navigation system [12]. With the cost of high-precision star sensor becoming lower and lower, the requirement of INS precision can be significantly reduced by introducing star sensor into the integrated. In [14], the attitude matrix output by star sensor is used as a measurement, which leads to the complexity of the filtering algorithm and the large amount of filtering calculation. According to the constraint conditions of attitude matrix, the diagonal elements are selected as one of the measurements of integrated navigation. The complexity of integrated navigation filtering algorithm is significantly reduced, and the amount of filtering calculation is reduced
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