• Journal of the Korea Institute of Military Science and Technology
• /
• v.13 no.3
• /
• pp.365-371
• /
• 2010

The fins of a missile which is folded within a canister are deployed according to a command during the missile flight. The aerodynamic load generated by operating environments such as missile flight speed, platform movement speed and wind acts as an anti-deploying force and prevents the fins from deploying. As the diversification of platforms and the higher speed of missiles need a larger deploying force but the space for operating the fin deploying device is getting narrower, the new design concepts are required for developing such a device. In this study, a pneumatic device for deploying missile fins is designed and its characteristics are verified through experiments and analyses.

• The Transactions of The Korean Institute of Electrical Engineers
• /
• v.63 no.1
• /
• pp.162-172
• /
• 2014

A three-dimensional dynamic model for simulating various motions of full vehicle is presented. The model has 16 independent degrees of freedom (DOF) consisting of three kinds of components; a vehicle body of 6 DOF, 4 independent suspensions equipped at every corner of the body, and 4 tire models linked with each suspension. The dynamic equations are represented in six coordinate frames such as world fixed coordinate, vehicle fixed coordinate, and four wheel fixed coordinate frames. Then these lead to the approximated prediction model of vehicle posture. Both lateral and longitudinal dynamics can be computed simultaneously under the conditions of which various inputs including steering command, driving torque, gravity, rolling resistance of tire, aerodynamic resistance, etc. are considered. It is shown through simulations that the proposed 3D model can be useful for precise design and performance analysis of any full vehicle control systems.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.03a
• /
• pp.97-104
• /
• 2008

Most of engine control systems for helicopter turboshaft engines are equipped with dual sensors. For the system with dual redundancy, analytic methods are used to detect faults based on the system dynamical model. Helicopter engine dynamics are affected by aerodynamic torque induced from the dynamics of the main rotor. In this paper an engine model including the rotor dynamics is constructed for the T700-GE-700 turboshaft engine powering UH-60 helicopter. The singular value decomposition(SVD) method is applied to the developed model in order to detect sensor faults. The SVD method which do not need an additional computation to generate residual uses the characteristics that the system outputs in direction of the left singular vector if an input is applied in direction of the right singular vector. Simulations show that the SVD method works well in detecting and isolating the sensor faults.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.03a
• /
• pp.686-689
• /
• 2008

The leaning angle optimization of turbine blade using the genetic algorithm was conducted in this paper. The calculation CFD technique was based upon the Diagonalized Alternating Directional Implicit scheme(DADI) with algebraic turbulence modeling. The leaning angle of VKI turbine blade was represented using B-spline curve. The control points are the design variable. Genetic algorithm was taken into account as an optimization tool. The objective was to minimize the total pressure loss. The optimized final geometry shows the better aerodynamic performance compared with the initial turbine blade.

• Advances in aircraft and spacecraft science
• /
• v.4 no.2
• /
• pp.93-124
• /
• 2017

Aeroelastic performance controls wing shape in flight and its behaviour under manoeuvre and gust loads. Controlling the wing‟s aeroelastic performance can therefore offer weight and fuel savings. In this paper, the rib orientation and the crenellated skin concept are used to control wing deformation under aerodynamic load. The impact of varying the rib/crenellation orientation, the crenellation width and thickness on the tip twist, tip displacement, natural frequencies, flutter speed and gust response are investigated. Various wind-off and wind-on loads are considered through Finite Element modelling and experiments, using wings manufactured through polyamide laser sintering. It is shown that it is possible to influence the aeroelastic behaviour using the rib and crenellation orientation, e.g., flutter speed increased by up to 14.2% and gust loads alleviated by up to 6.4%. A reasonable comparison between numerical and experimental results was found.

• Journal of the Korea Institute of Military Science and Technology
• /
• v.19 no.4
• /
• pp.476-482
• /
• 2016

In this study, a method using single Wheatstone bridge flexure has been presented to measure hinge moment acting on control surfaces of wind tunnel models. The structural simplicity of the flexure reduces difficulty regarding gauging and wire-routing, and also makes it feasible to install flexures even inside thin wings. Some flexures were designed and fabricated under typical aerodynamic loads in wind tunnel test, and the strains on the flexure according to applied loads were compared with the result of the analysis by finite element method. The relation between applied loads and output signals showed good linearity, and the standard deviation on the residual errors from linear equation obtained by least square method was within 1.0 % of the maximum design moments. In addition, the FEM analysis on the thickness of load-connecting part of the flexure showed that the sensitivity was improved as the thickness became thin as much as desired to avoid buckling.

• Smart Structures and Systems
• /
• v.3 no.1
• /
• pp.89-114
• /
• 2007

Two aspects of the design of a small-scale smart wing are addressed in this work, related to the ability of the wing to modify its cross section assuming the shape of two different airfoils and to the possibility of deflecting the profiles near the trailing edge in order to obtain hingeless control surfaces. The actuation is provided by one-way shape memory alloy wires eventually coupled to springs, Shape Memory Alloys (SMAs) being among the most promising materials for this kind of applications. The points to be actuated along the profiles and the displacements to be imposed are selecetd so that they satisfactorily approximate the change from an airfoil to the other and to result in an adequate deflection of the control surface; the actuators and their performances are designed so that an adequate wing stiffness is guaranteed, in order to prevent excessive deformations and undesired airfoil shape variations due to aerodynamic loads. The effect of the pressure distributions, calculated by way of the XFOIL software, and of the actuators loads, is estimated by FE analyses of the loaded wing. Two prototypes are then realised incorporating the variable airfoil and the hingeless aileron features respectively, and the verification of their shapes in both the actuated and non-actuated states, supported by image analysis techniques, confirms that interesting results are achievable with the proposed lay out and design considerations.

• Journal of Institute of Control, Robotics and Systems
• /
• v.11 no.1
• /
• pp.84-91
• /
• 2005

In this paper, we describe the real-time parallel processing simulator developed for the use of performance analysis of rolling missiles. The real-time parallel processing simulator developed here consists of seeker emulator generating infrared image signal on aircraft, real-time computer, host computer, system unit, and actual equipments such as auto-pilot processor and seeker processor. Software is developed from mathematic models, 6 degree-of-freedom module, aerodynamic module which are resided in real-time computer, and graphic user interface program resided in host computer. The real-time computer consists of six TIC-40 processors connected in parallel. The seeker emulator is designed by using analog circuits coupled with mechanical equipments. The system unit provides interface function to match impedance between the components and processes very small electrical signals. Also real launch unit of missiles is interfaced to simulator through system unit. In order to apply the real-time parallel processing simulator to performance analysis equipment of rolling missiles it is essential to perform the performance verification test of simulator.

• International Journal of Aeronautical and Space Sciences
• /
• v.16 no.3
• /
• pp.339-346
• /
• 2015

Aerodynamic flow control phenomena were investigated with a low-current DC surface discharge plasma actuator. The plasma actuator was found to operate in three different discharge modes with similar discharge currents of about 1 mA or less. Stable continuous DC discharge without audible noise was obtained at higher ballast resistances and lower discharge currents. However, even with continuous DC power input, a low-frequency self-pulsed discharge was obtained at lower ballast resistances, and a high-frequency self-pulsed discharge was obtained at higher set-point currents and higher ballast resistances, both with audible noise. The Schlieren image reveals that the low-frequency self-pulsed mode produces a synthetic jet-like flow implying that a gas heating effect plays a role, even though the discharge current is small. The high-frequency self-pulsed mode produces pulsed jets in a tangent direction, and the continuous DC mode produces a steady straight pressure wave. Particle image velocimetry (PIV) images reveal that the induced flow field by the low-frequency self-pulsed mode has flow propagating in the radial direction and centered between the electrodes. The high-frequency self-pulsed mode and continuous DC mode produce flow from the anode to the cathode. The perturbed region downstream of the cathode is larger in the high-frequency self-pulsed mode with similar maximum speeds.

• Journal of the Korean Society for Aviation and Aeronautics
• /
• v.23 no.4
• /
• pp.22-29
• /
• 2015

Quadrotor is widely used in variable application nowadays. Due to its inherent unstable characteristics, control system to augment the stability is essential for quadrotor operation. To design control system and verify its performance through simulation, accurate dynamic model is required. Quadrotor dynamic model is simply compared with conventional rotorcraft such as helicopter. However, the accurate dynamic model of quadrotor is not easy to develop because of the highly correlated aerodynamic effect of each rotor. In this paper, quadrotor dynamic model is identified from the flight data using frequency domain approach. Flight test of quadrotor is performed in closed loop configuration with stability augmentation system included. Frequency sweep input is applied in each of lateral, longitudinal, yaw and heave axis separately. The bare dynamic model is identified from the flight data of quadrotor responses and thrust measurement through Pulse Width Modulation(PWM) data. The frequency responses of identified model match well with those of flight data, and time responses of identified model for doublet input in each axis are also shown to agree with flight data.