In order to predict the result of impact test in the design phase and reduce the experimental times, which can save cost and shorten development cycle, a finite element model of aluminum alloy wheel 13-degree impact test is established based on Abaqus. All mechanical parts such as the standard impact block, the assembly of the wheel and the tire, the support and bolts are included in the finite element model. The predicted result of finite element analysis and the experimental result agree very well shows the finite element model is correct. The equivalent plastic strain value was also put forward as fracture criterion for the wheel in the impact test which realizes the transition from the qualitative analysis to the quantitative analysis in the development process of aluminum alloy wheel.
There is the long periodicity attitude error between true attitude and measurement attitude using star sensor for spacecraft attitude determination system because of aberration of light. Aberration of light occurs because the spacecraft’s velocity has a component that is perpendicular to the line traveled by the light incoming from the star. The type of aberration is analyzed and their constants of aberration are calculated in this paper. According to the constants the aberration, the correction mathematical models of parallax of aberration of light of these types of aberration are derived. The parallax of aberration of light of the recognized stars in the FOV of star sensor is calculated with the mathematical models. Then the true vectors of recognized stars at image space coordinate system of star sensor are calculated. The measurement attitude of star sensor is calculated with the true vectors of recognized stars and their vectors at celestial sphere coordinate system. The simulations show the long periodicity attitude error is corrected with the method in this paper. At last the correction of aberration of light was successfully demonstrated using two star sensors with real sky experiment in 2011.
Aiming at the limitations of the orbital dynamic equations based star sensor navigation method, a star sensor /geomagnetic information utilized aircraft autonomous navigation method is proposed. Dynamic equations applicable to general aircrafts are established. System observation equations are deduced. The angle between geomagnetic and starlight vector is used as observation in the algorithm. Extended Kalman filter is used to estimate position and velocity of aircraft in the algorithm. Singular value decomposition method is used to analyze observability of the system. Simulation results show that the algorithm has many advantages including high precision, good filtering convergence and stability, and non-accumulated error. The algorithm can be used as aided navigation of inertial navigation or in occasions, which only require a general navigation precision.
The larger area threshold scan window in star image is scanned with conventional star tracking algorithm under large maneuvering of the vehicle. This results that the number of calculating angular separation between two stars is increased. So the update rate of star sensor is decreased and the feasibility of errors star locations is increased. So, an autonomous predicted star locations algorithm with high angle velocity is presented in this paper. Firstly, the actual star locations are obtained within the threshold scan window of corresponding ideal star locations calculated by last frame in the current star image. Secondly the current attitude information is calculated with the obtaining actual star locations. Thirdly, the next ideal attitude is calculated with current attitude and last attitude. Finally, the next ideal star locations are calculated with the next ideal attitude. At last the algorithm are tested not only simulation but also at night sky experiment with the angular speed maneuvering of aircraft is 5.208681°/s in 2013. And the algorithm has be applied star sensor of satellite GNC.
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