• 제어로봇시스템학회:학술대회논문집
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• 2003.10a
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• pp.1211-1215
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• 2003

Accurate mathematical models of spacecraft components are an essential of spacecraft attitude control system design, analysis and simulation. Gyro is one of the most important spacecraft components used for attitude propagation and control. Gyro errors may seriously degrade the accuracy of the calculated spacecraft angular rate and of attitude estimates due to inherent drift and bias errors. In order to validate this model, nominal case simulation has been performed and compared for the low range mode and high range mode, respectively. In this paper, a mathematical model of gyro containing the relationships for predicting spacecraft angular rate and disturbances is proposed.

• International Journal of Aeronautical and Space Sciences
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• v.11 no.3
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• pp.234-239
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• 2010

The utilization of micro-electro-mechanical system (MEMS) gyros and accelerometers in low-level inertial measurement unit (IMU) influences cost effectiveness in a positive way under the condition that device error characteristics are fully calibrated. The conventional calibration process utilizes a rate table; however, this paper proposes a new method for achieving reference calibration data from the natural motion of pendulum to which the IMU undergoing calibration is attached. This concept was validated with experimental data. The pendulum angle measurements correlate extremely well with the solutions acquired from the pendulum equation of motion. The calibration data were computed using the regression method. The whole process was validated by comparing the measurement from the 6 sensor components with the measurements reconstructed using the identified calibration data.

• 제어로봇시스템학회:학술대회논문집
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• 2001.10a
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• pp.167.4-167
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• 2001

Accurate mathematical models of spacecraft components are an essential of spacecraft attitude control system design, analysis and simulation. Gyro is one of the most important spacecraft components used for attitude propagation and control. Gyro errors may seriously degrade the accuracy of the calculated spacecraft angular rate and of attitude estimates due to inherent drift and bias errors. In this paper, a detailed mathematical model of gyro containing the relationships for predicting spacecraft angular rate and disturbances is proposed. Stochastic model describing random drift behavior is discussed in frequency domain and time domain. In order to illustrate this approach, we analyze the behavior for Litton´s Space Inertial Reference Uint(SIRU).

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.4
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• pp.308-315
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• 2016

A new technique for Attitude & Heading Reference System (AHRS) by using sequential measurement noise covariance (SMNC) is addressed in a vibration environments in this paper. In particular, a low-cost inertial measurement unit in general diverges in the acceleration phase or vibrating environments due to inherent properties of gravity and acceleration. In this paper, by considering current and prior measurements to estimate actual attitudes and headings in a local frame, the proposed technique overcomes these problems efficiently. Finally, the performance of the suggested approach is verified by numerical simulations.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.50 no.9
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• pp.617-625
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• 2022

This study investigates the use of movable calibration target for gyroscopic and accelerometer bias compensation of inertial measurement units for firing lane alignment. Calibration source is detected with the help of vision sensor and its information in fused with other sensors on launcher for error correction. An algorithm is proposed and tested in simulation. It has been shown that it is possible to compensate sensor biases in firing launcher in few seconds by accurately estimating the location of calibration target in inertial frame of reference.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.9
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• pp.734-745
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• 2017

The navigation system of lunar lander are composed of various navigation sensors which have a complementary characteristics such as inertial measurement unit, star tracker, altimeter, velocimeter, and camera for terrain relative navigation to achieve the precision and autonomous navigation capability. The required performance of sensors has to be determined according to the landing scenario and mission requirement. In this paper, the specifications of navigation sensors are investigated through covariance analysis. The reference error model with 77 state vector and measurement model are derived for covariance analysis. The mission requirement is categorized as precision exploration with 90m($3{\sigma}$ ) landing accuracy and area exploration with 6km($3{\sigma}$ ), and the landing scenario is divided into PDI(Powered descent initiation) and DOI(Deorbit initiation) scenario according to the beginning of autonomous navigation. The required specifications of the navigation sensors are derived by analyzing the performance according to the sensor combination and landing scenario.