• Journal of the Institute of Electronics and Information Engineers
• /
• v.54 no.7
• /
• pp.127-132
• /
• 2017

In this paper, we propose a controller gain based on model based design and implement the two-wheel inverted pendulum type robot using NXT Lego and RobotC language. Two-wheel inverted pendulum robot consists of NXT mindstorm, servo DC motor with encoder, gyro sensor, and accelerometer sensor. We measurement wheel angle using bulit-in encoder and calculate wheel angle speed using moving average method. Gyro measures body angular velocity and accelerometer measures body pitch angle. We calculate body angle with complementary filter using gyro and accelerometer sensor. The control gain is a weighted value for wheel angle, wheel angular velocity, body pitch angle, and body pich angular velocity, respectively. We experiment and observe the effect of two-wheel inverted pendulum with respect to change of control gains.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.37 no.10
• /
• pp.992-997
• /
• 2009

This paper focuses on the calculation of the missile angular velocity under the reduced sensor condition and its verification using the Flight Motion Simulator(FMS). The missile angular velocity is usually measured by the body gyroscopes, but we assume that the inertial sensors on the missile body are in the absence of pitch and yaw gyroscopes. Under this reduced sensor condition, this paper shows the missile angular velocity can be calculated by using the gimbal seeker gyroscope, the roll body gyroscope, the gimbal angle and its rate. The FMS experiment was carried out to verify the proposed algorithm.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.34 no.5
• /
• pp.65-73
• /
• 2006

When a satellite separates from the launch vehicle, an initial high angular rate or a tip-off rate is generated. B-dot logic is generally used for controlling the initial tip-off rate. However, it has the disadvantage of taking a relatively long time to control the initial tip-off rate. To solve this problem, this paper suggests a new detumbling control method to be able to adapt to micro/nanosatellite with the pitch bias momentum system. Proposed detumbling method was able to control the angular rate within 20 minutes which is significantly reduced compared to conventional methods. Since the previous wheel start-up method cannot be used if the detumbling controller proposed by this paper is used, a method is also proposed for bringing up the momentum wheel speed to nominal rpm while maintaining stability in this paper. The performance of the method is compared and verified through simulation. The overall result shows much faster control time compared to the conventional methods, and achievement of the nominal wheel speed and 3-axes stabilization while maintaining stability.

• Journal of the Institute of Convergence Signal Processing
• /
• v.9 no.2
• /
• pp.159-163
• /
• 2008

In this paper, a simplified type of adaptive controller using Nussbaum gain for the control of the spacecrapt with elastic appendages is suggested. This method doesn't need the information of the high frequency components in transfer function. While the pitch angle tracks the desired value by this method, the elastic modes are also stabilized. Only pitch angle and the pitch rate are used for the design of the output feedback controller. Especially all system parameters and the high frequency gain are assumed to be unknown. For design simplicity, a controller is designed by using only the linear part, and it's shown to satisfy the nonlinear system by the simulation with basic explanations. By using the Lyapunov function, the stability of the suggested algorithm is demonstrated, and also the effectiveness of the suggested algorithm is verified by showing the computer simulation results.

• Journal of the Korea Institute of Military Science and Technology
• /
• v.5 no.2
• /
• pp.37-49
• /
• 2002

In this paper, flyout stability of missile that is launched in inclined launcher using sabots is analyzed. To include missile bending motion during flyout, FEA model of missile is converted into eight concentrated mass and equivalent stiffness matrix. Six d.o.f ship motion that have influence on flyout stability is modeled and missile firing time is modeled as probability variable to take arbitrary ship attitude into account. Gap between missile and sabot is modeled as normal distribution probability variable and Monte Carlo simulation is performed. As results, the coriolis acceleration effects by ship motion are analyed and statistical results of missile pitch rate are shown.

• Journal of the Korea Institute of Military Science and Technology
• /
• v.2 no.2
• /
• pp.92-98
• /
• 1999

Dynamic behavior of missile which is fired in canister is analyzed by flexible multibody dynamics. The bending elasticity of missile is very important in case that missile is fired in the inclined launcher. In this paper, the force element model for the missile launching stage and the finite element model of missile are developed. The FEA model of missile is condensed into five lumped mass element model and the consistence between FE model and lumped mass model of missile is verified by modal analysis. As a result of analysis, sabot reaction force and pitch rate of missile for the variation of gap size and force element are obtained.

• Journal of Advanced Marine Engineering and Technology
• /
• v.36 no.5
• /
• pp.635-644
• /
• 2012

In order to design a wave energy generating system, a 6-DOF analysis technique is applied to the three-Dimensional CFD analysis on of a floating body and the behavior is interpreted according to the nature of the incoming wave. A wave period of 5.5s & amplitude of 0.57m from Marado is chosen. 12 case of natural pitching period from 1.25 to 2.8s has been modeled. The relation between tuning factor & pitch angle for the waves generated is compared to analyze the effects of energy absorption variables, namely mass moment of inertia, angular velocity and angular acceleration. From the results obtained, we conclude that model L is the maximum power absorbed, 6kW approximately. A maximum pitch angle of 1.91 degree was attained by Model F, and the maximum displacement of nearly 0.7m was attained by Model L among models D, F and L.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.39 no.2
• /
• pp.162-169
• /
• 2011

The flight control system utilize RSS(Relaxed Static Stability) criteria in both longitudinal axes to achieve performance enhancements and improve stability. The aircraft using digital flight-by-wire flight control system receives aircraft flight conditions such as pitch, roll and yaw rate, normal acceleration from RSA(Rate Sensor Assembly) and ASA(Acceleration Sensor Assembly). These sensors has permissible measurement error related to system safety of an aircraft but, unexpected flight motions are happened by sensing errors such as offset, noise and etc. The unexpected pitch down tendency occurred by ASA sensor bias in 1g level flight with pilot hands-off. This paper addresses the design and verification of flight control law to improve of pitch down or up tendency caused by ASA sensor bias. The result of analysis and flight test reveals that pitch down tendency can be improved by pitch attitude feedback system.

• Bulletin of the Korean Space Science Society
• /
• 2006.04a
• /
• pp.121-124
• /
• 2006

B-dot logic is generally used for controlling the initial tip-off rate. However, it has the disadvantage of taking a relatively long time to control the initial tip-off rate. To solve this problem, this paper suggests a new detumbling control method to be able to adapt to micro/nanosatellite with the pitch bias momentum system. Proposed detumbling method was able to control the angular rate within 20 minutes which is a significant reduction compared to conventional methods.

• Journal of the Institute of Electronics and Information Engineers
• /
• v.53 no.7
• /
• pp.95-102
• /
• 2016

Geomagnetic sensors are widely utilized for sensing heading direction of quadrotors. However, measurement from a geomagnetic sensor is easily corrupted by environmental magnetic field interference and roll/pitch directional motion. In this paper, a measurement method of a quadrotor heading direction is proposed for application to yaw attitude control. In order to eliminate roll/pitch directional motion effect, the geomagnetic sensor data is compensated using the roll/pitch angles measured for stabilization control. In addition, yaw-directional angular velocity data from a gyroscope sensor is fused with the geomagnetic sensor data using a complementary filter which is a simple and intuitive sensor fusion method. The proposed method is applied to experiments, and the results are presented to prove validity and effectiveness of the proposed method.