Initial Attitude Determination Research Articles

측정치 시간지연을 가진 전자광학 추적장비의 초기 자세 결정 알고리즘

In order to have accurate target positions of the EOTS(Electro-Optical Tracking System), more accurate navigation outputs are required. An IMU(Inertial Measurement Unit) can be attached to the EOTS to have high rate INS ouputs. Due to the time-delay of the vehicle’s navigation outputs, it is not easy to have accurate target positions. When the vehicle moves at high speed and with high maneuverability, the accuracy becomes even worse. In this paper, an initial attitude determination algorithm of the EOTS with measurement time-delay is proposed. In order to compensate the measurement time-delay, the time-delay variable is augmented into the integrated Kalman filter for the IMU of EOTS and vehicle’s measurement output. The proposed algorithm was evaluated for sensor’s model-based measurement generator outputs, and a land vehicle experiments were performed. The evaluation results show that the proposed algorithm compensates the measurement time-delay well and gives more accurate initial attitude.

Initial Attitude Determination of Aerial Platform Based on MIMU

Aiming at the problem of initial attitude determination of aerial platform, a low computational attitude fusion algorithm is proposed based on MIMU. Attitude kinematic equations described by quaternions are developed and the filter structure used in this paper is designed. Real fly test results show that: at different initial deviation, the alignment precision of roll and pitch angles are within 1 °, attitude errors converge to 5 °or less within 12s, which meets the requirements of the aerial platform attitude fast self-determination. The filter algorithm has a small amount of calculation and is easy to be realized.

An Integrated GPS/INS Navigation System for Land Vehicle

A precise guided system needs an efficient control depending on a precise navigation algorithm, with the ability of getting an accurate initial attitude determination to guarantee the mission success. A navigation system is presented in this paper based on integration between inertial measuring unit and Global Positioning System via Kalman filter approach to satisfy an acceptant accuracy. The two well known Euler and Quaternion attitude determination techniques are implemented to evaluate the body orientation during motion. The carried out system is validated using both simulation data and experimental work. The simulation data is obtained using a six-degree-of-freedom model for a 122mm artillery rocket to obtain all ballistic trajectory parameters during flight. The experimental work is done using a land vehicle taking into consideration the initial attitude determination problem. The results showed high accuracy improvements with high data rates 200 Hz for full state navigation information (position, velocity and attitude).

B8 航空機実験によるGPS姿勢決定アルゴリズム評価

Spaceborne GPS navigation technology is recently studied very actively because it is effective to enhance the autonomy of space vehicles on Low Earth Orbit (LEO). GPS attitude determination uses carrier phase measurements observed at multiple antennas, and especially it contributes to simplification of the GN&C sub-system. The major task of GPS attitude determination is to improve reliability and accuracy. Especially an initial attitude search algorithm, which resolves integer ambiguity of carrier wave length, should not be affected by measurement error. In this paper, we show the algorithm for initial attitude determination by integer search technique. Residuals to be evaluated for initial search are accumulated in order to mitigate the effects of measurement noise. It was proved to be effective by ground field tests. And we evaluated it by aircraft flight experiments, which were conducted in November 2001. Ambiguity solution and attitude can be obtained under dynamic attitude maneuvers.

Initial attitude determination for the hipparcos satellite

The present paper described the strategy and algorithms used during the initial on-ground three-axes attitude determination of ESA's astrometry satellite HIPPARCOS. The estimation is performed using calculated crossing times of identified stars over the Star Mapper's vertical and inclined slit systems as well as outputs from a set of rate-integrating gyros. Valid star transits in either of the two fields of view are expected to occur in average about every 30 s whereas the gyros are sampled at about 1 Hz. The state vector to be estimated consists of the three angles, three rates and three gyro drift rate components. Simulations have shown that convergence of the estimator is established within about 10 min and that the accuracies achieved are in the order of a few arcsec for the angles and a few milliarcsec per s for the rates. These stringent accuracies are in fact required for initialisation of subsequent autonomous on-board real-time attitude determination.