• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.6
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• pp.506-514
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• 2012

This paper represents a method for designing of path following Line-of-Sight(LOS) guidance law based on vehicle kinematics. In general, a LOS guidance law which is composed of gains and approach length as design parameters is designed by empirical or trial-and-error method. These approaches cannot guarantee a precision tracking performance of guidance law consistently. Also, the design parameters should be redesigned with variations of vehicle maneuverability and flight velocity. Based on a vehicle kinematics with its velocity, the proposed method for designing of parameters not only minimizes the number of design parameters, also has a reliable and consistent tracking performance using variable guidance gain changed in accordance with flight velocity. This is validated by nonlinear simulation with $1^{st}$ order attitude response dynamics and flight experiments with given linear and circular path.

• Journal of Advanced Navigation Technology
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• v.21 no.5
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• pp.450-458
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• 2017

This paper presents dynamic modelling and simulation study on attitude/altitude control of an insect-mimicking flapping micro aerial vehicle during hovering. Mathematical modelling consists of three parts: simplified flapping kinematics, flapping-wing aerodynamics, and six degree of freedom dynamics. Attitude stabilization is accomplished through linear quadratic regulator based on the linearized model of the time-varying nonlinear system, and altitude control is designed in the outer loop using PID control. The performance of the proposed controller is verified through numerical simulation where attitude stabilization and altitude control is done for hovering. In addition, it is confirmed that the attitude channel by periodic control is marginally stable against periodic pitching moment caused by flapping.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.7
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• pp.600-609
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• 2017

We designed a target management integration system that enables us to select the final target manually or automatically from seeker's sensor image. The integrated system was developed separately for the air vehicle system and the ground system. The air vehicle system simulates the motion dynamics and the sensor image of the air vehicle, and the ground system is composed of the target template image module and the ground control center module. The flight maneuver of the air vehicle is based on pseudo 6-degree of freedom motion equation and the proportional navigation guidance. The sensor image module was developed using the known infrared(IR) image rendering method, and was verified by comparing the rendered image to that of a commercial software. The ground control center module includes an user interface that can display as much information to meet user needs. Finally, we verified the integrated system with simulated impact target mission of the air vehicle, by confirming the final target change and the shot down result of the user's intervention.

• Journal of the Korea Society for Simulation
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• v.29 no.3
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• pp.49-55
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• 2020

A goal of improving projectile is to increasing achievable range. The shape of a projectile is generally selected on the basis of combined aerodynamics and structural considerations. The choice of body, nose and boattail shape has a large effect on aerodynamic design. One of the main design factors that affect projectile configuration is aerodynamic drag. The aerodynamic drag refers to the aerodynamic force that acts opposite to the relative motion of a projectile. An investigation was made to predict the effects of nose, boattail and body shapes on the aerodynamic characteristics of projectiles using a semi-empirical technique. A parametric study is conducted which includes different projectile geometry. Performance predictions of achievable range are conducted using a trajectory simulation model. The potential of extending the range of a projectile using optimization of projectile configuration is evaluated. The maximum range increase is achieved due to the combination of optimal body shapes.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.1
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• pp.42-54
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• 2009

Contrast to the 6-DOF nonlinear dynamic modeling of nonlinear tracking problem, 3-DOF point-mass modeling of flight mechanics is efficient and adequate for applying the trajectory optimization problem. There exist limitations to apply an optimal trajectory from point-mass modeling as a reference trajectory directly to conduct the nonlinear tracking control, In this paper, new matching trajectory optimization scheme is proposed to compensate those differences of mismatching. To verify performance of proposed method, full ascent three-dimensional flight trajectories are obtained by reflecting the real constraints of flight conditions and airship performance with and without jet stream condition. Then, they are compared with the optimal trajectories obtained from conventional method.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.7
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• pp.647-654
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• 2010

In this study, an efficient ornithopter aerodynamic model, which is applicable to ornithopter wing design considering fluid-structure interaction or ornithopter flight dynamics and control simulation, was proposed and experimentally validated through the wind tunnel experiments. Due to the ornithopter aerodynamics governed by unsteady low Reynolds number flow, an experimental device was specially designed and developed. A part of the experimental device, 2-axis loadcell, was situated in the non-inertial frame; the dynamic calibration method was established to compensate the inertial load for pure aerodynamic load measurements. The characteristics of proposed aerodynamic model were compared with the experimental data in terms of mean and root-mean-square values of lift and drag coefficients with respect to the flow speed, flapping frequency, and fixed angle of attack.

• Journal of the Institute of Electronics Engineers of Korea SC
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• v.46 no.1
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• pp.87-96
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• 2009

In conventional method, flight model is discribed to differential equation by linealization of nonlinear object motion equation. As state equation from differential equation of moving object, the controller is designed by transfer functions of each module under discrimination of stability criteria. But this conventional method is designed under limitation of nonlinearity from object's shape and speed. In other word, The greater part of guidance/navigation system was satisfied with the result of good performance for normal figure of flight object, not sudden changed flight condition, not high speed. But it is not able to give full play to its ability on flight object which has abnormal figure, sudden changeable motion, high speed. Therefore, in this paper was presented performance analysis of load control model for navigation/guidance system on flying object being uncertainty, non-linear like abnormal figure, sudden changeable motion, high speed and is presented method of trajectory control(controllability) ahead of controllability and stability to achieve flight mission. In other word, this paper shows the first step of Min-design method and flight control model.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.6
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• pp.92-100
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• 2002

In this study, transonic/supersonic nonlinear flutter analysis system of a complete aircraft including forced harmonic motion pf control surfaces has been effectively developed using the modified transonic small disturbance (TSD) equation. To consider the nonlinear effects, the coupled time marching method (CTM) combining computational structural dynamics (CFD) has been directly applied for aeroelastic computations. The grid system for a complex full aircraft configuration is effectively generated by the developed inhouse code. Intransonic and supersonic flight regimes, the characteristics of static and dynamic aeroelastic effect has been investigated for a complete aircraft model. Also, nonlinear flutter flight simulations for the forced harmonic motion of control surfaces are practically presented in detail.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.10
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• pp.905-911
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• 2011

With the progress of micro actuator technology, studies on the development of micro air flapping wing vehicles are actively undergoing. In the present study, the changes of both lift and thrust characteristics of the wings are investigated using a boundary element method. Lift of the heaving wing is not generated when the wing is beating with smaller frequencies than 1 Hz. Thrust increases with amplitude and frequency. As the wing's taper and aspect ratios increase, both lift and thrust also increase. Results on the pitching oscillation and flapping motion will be included in the future work.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.5
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• pp.395-404
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• 2012

This paper presents computations of the dynamic derivatives of three dimensional flight vehicle configurations using CFD. The pitch dynamic derivatives are computed from the pitch sinusoidal motion, while the roll damping is computed based on steady state calculation using a non-inertial frame method. The Basic Finner and the SDM(Standard Dynamic Model) are chosen for the benchmark tests against other numerical and experimental results. For the flow calculations, a 3-D Euler solver that can be run both on the non-inertial frame and on the inertial frame is developed. A dual-time stepping method is applied for the unsteady time accurate simulations. A good agreement of pitch-roll dynamic derivatives with previously published numerical results and the experimental results is observed.