Integrated navigation method based on star sensor calibration

Abstract

The combined navigation based on star sensor calibration optimizing the inertia, astronomical and satellite for all kinds of high precision fly guide. It modifies the installation error based on the star diagonal distance to get high precision information, calculating the disguise distance and disguise distance ratio relative to the inertia guide position through the astronomical data provided by GPS and the location and speed provided by the inertia guide and comparing with the disguise distance and its ratio measured by the GPS as the measuring value, through combined Kalman filter model to estimate the inertia guide system and GPS error and ratifying accordingly to these two systems based on feedbacks. It can improve combined guide navigation precision and reach a high quality combination, whose precision is better than any single guide navigation precision.

Description

technical field [0001] The invention relates to an inertial / astronomy / satellite combined navigation method, in particular to a deep integrated combined navigation method based on star sensor calibration, which can be used for various high-precision aircraft navigation. Background technique [0002] Existing various single navigation systems all have their advantages and disadvantages. The inertial navigation system is a completely autonomous positioning and navigation system. It can continuously provide position, speed and attitude information in real time. Its short-term accuracy is very high, but the error of the inertial system will continue to accumulate over time. Therefore, pure inertial guidance The system is difficult to meet the requirements of long-range high-precision navigation. The satellite navigation system can provide the three-dimensional position and speed information of the carrier in real time all-weather, and the error does not accumulate over time, and...

AI Technical Summary

Problems solved by technology

But its disadvantage is that it cannot provide high-precision attitude information, the navigation information is discontinuous, and it is easy to be interfered.

Usually the integrated navigation system adopts a simple combination method, using the error equation of the inertial navigation system as the state equation of the integrated navigation system, taking the position, velocity and attitude information output by satellite navigation and celestial navigation as observations, and using Kalman filter to combine Navigation information processing corrects inertial navigation errors. The biggest advantage of this combination method is that it is easy to implement and improves the navigation accuracy of the system. However, the accuracy of integrated navigation depends on the accuracy of observations. The accuracy of position and speed is equivalent to that of GPS, and the accuracy of attitude is comparable to The attitude accuracy determined by the star sensor is equivalent, but the optimal combination of each navigation system cannot be achieved, and the accuracy of the integrated navigation system is better than that of any single navigation system, and the error accumulation caused by the model over time and some system errors Any sudden change will cause error transmission, which often leads to the instability of the combined filter. Therefore, the error of the auxiliary navigation system must be further compensated to obtain higher navigation accuracy.

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