SINS Guidance Instrument Error Compensation

With the developnent of computer technology and the guidance technology, SINS guidance instrument errors become the key factors to influence the homing precision of missile. Limited by the processing and manufacturing level, however, the improvement of inertial navigation system accuracy merely by increasing the machining accuracy of inertial devices such as gyroscopes and accelerometers has become more and more costly. By contrast , a more economic method is the system level or the instrument level guidance instrument compensation based on the inertial instruments. The effectiveness of compensation depends on the accuracy of inertial instrument error model, the stability of the error coefficients and the perfection of compensation method. Therefore, many problems in SINS guidance instrument error separation and compensation are deeply and systematically investigated from both theory and practical aspects. A series of research work have been done in the respects of experimental design, model structure identification, identification algorithm, consistence check between the flight and the ground, error compensation.Firstly, the SINS guidance instrument error model is built. Static and dynamic error models of inertial sensors are deduced from the types of interferential torque and dynamic principle. The SINS guidance instrument error model is deduced based on the principle of SINS. The physical meaning of error model is very explicit. And the dimension of it is easy to be augmented or reduced according to the need of simulation study or engineering application.Then the algorithm of SINS in high dynamic circumstance is deeply investigated in order to provide the high-precision velocity error between the telemetric velocity and tracking velocity as the observation of the SINS guidance instrument error model. The scrolling compensation in position solution, sculling compensation in velocity solution and the coning compensation in attitude solution are emphasized so that the algorithmic error is decreased. On the other hand, because of false-negative judgments and miss judgments of outlier values, the fuzzy forecasting system based on fuzzy cluster analysis and neural network is brought out to detect and delete the abnormal values. The ballistic is reconstructed accurately based on the pulse outputs of inertial instruments. Next, considering the characteristics of high grades and serious multiliearity of SINS guidance instrument error model, the scheme of structure identification first and parameters identification later is applied. The common method, that all strong-correlation terms of the model are eliminated, can bring the loss in the engineering application, so the new method is proposed that the identified model reserves some correlation. The augmented matrix A is constructed by the outputΔW and the matrix S. The"determinating order based on ratio of determinant"is brought out to screen the strong-correlation terms in the structure identification. The latent root estimation is improved in screening the eigenvalues and eigenvectors. Thus the estimation precision is improved greatly.The consistence check of guidance instrument error coefficients of flight test and ground test is the purpose of flight experiment. The causes of inconsistency of the two models are analyzed. The hypothesis test of linear regression model based on F statistics is proposed to check the consistence.Finally, the instability of error coefficients is probably caused by the change of the flight environments, therefore, the relation between the error coefficients and flight environment is analyzed. The approach is presented to identify SINS guidance instrument error models and compensate the error in the segmented sections corresponding to the change of vertical acceleration of aircraft. It can decrease the multiliearity of model and computational complexity greatly so that the identification accuracy of error model and guidance precision is improved greatly.