• Journal of the Korea Institute of Military Science and Technology
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• v.10 no.1
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• pp.33-43
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• 2007

Using the more common conventional chordwise aerodynamic approach, flapping a flat plate wing with zero degree chordwise pitch angle of attack and no relative wind should not produce lift. However, in hover, with no forward relative velocity and zero degree chordwise pitch angle of attack, flapping flat plate wings does in fact produce lift. In the experiments peformed for this paper, the flapping motion is considered pure(downstroke and upstroke) with no flapping stroke plane inclination angle. No changes in chordwise pitch angle are made. The total force is measured using a force transducer and the net aerodynamic force is determined from this measured total force by subtracting the experimentally determined inertial contribution. These experiments were repeated at various flapping frequencies and for various wing planform sizes for flat plate wings. The trends in the aerodynamic lift variation found using a force transducer have nearly identical shape for various flapping frequencies and wing planform sizes.

• Journal of Computing Science and Engineering
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• v.7 no.2
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• pp.122-131
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• 2013

The MPI CyberMotion Simulator provides a unique motion platform, as it features an anthropomorphic robot with a large workspace, combined with an actuated cabin and a linear track for lateral movement. This paper introduces the simulator as a tool for studying human perception, and compares its characteristics to conventional Stewart platforms. Furthermore, an experimental evaluation is presented in which multimodal human control behavior is studied by identifying the visual and vestibular responses of participants in a roll-lateral helicopter hover task. The results show that the simulator motion allows participants to increase tracking performance by changing their control strategy, shifting from reliance on visual error perception to reliance on simulator motion cues. The MPI CyberMotion Simulator has proven to be a state-of-the-art motion simulator for psychophysical research to study humans with various experimental paradigms, ranging from passive perception experiments to active control tasks, such as driving a car or flying a helicopter.

• Journal of computational fluids engineering
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• v.16 no.4
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• pp.21-27
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• 2011

In this paper, the diagonal implicit harmonic balance method is applied to analyze helicopter rotor blade flow. The periodic boundary condition for Fourier coefficients is also applied in hover and forward flight conditions. It is available enough to simulate the forward flight problem with only one rotor blade using the periodic boundary condition in the frequency domain. In order to demonstrate the present method, Caradonna & Tung's rotor blades were used and the results were compared to the time-accurate method and experimental data.

• Journal of computational fluids engineering
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• v.17 no.2
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• pp.78-84
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• 2012

Numerical calculations were performed to investigate the influence in aerodynamic characteristics of a rotor system by surrounding structures and the ground effect for the rotor blade on a whirl tower is also investigated. Three dimensional Navier-Stokes simulations were carried out by using unstructured overset mesh technique and parallel computation. The calculated hover performance showed good agreement with the experimental result and showed that the structures around the whirl tower did not affect the aerodynamic characteristics of the blade. The ground effect was studied by comparing with the numerical result for the out of ground condition and the result of an analytic model.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 1998.04a
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• pp.386-391
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• 1998

The effect of the changes in parameter angles(precone, droop, sweep) on the lead-lag damping was focused on. Experiment was made with hingeless 4-blade rotors and NACA 0012 airfoil. For the measurement of the rotating natural frequencies and lead-lag damping, non-rotating swash plate was oscillated at the regressing lag mode frequency and the data were acquired after the excitation was cut off. Analysis was made using a finite element formulation based on Hamilton's principle. The main blade is assumed as elastic beams. Quasi-steady strip theory is used to obtain aerodynamic forces, and non-circulatory forces are also included.

• 한국항공운항학회:학술대회논문집
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• 2016.05a
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• pp.9-13
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• 2016

Reaction drive rotor system is capable of providing hover and low speed capabilities to different aircraft concepts such as stopped rotor wing, canard rotor wing, compound gyroplane etc. Existing sizing and analysis tools for shaft drive rotor system cannot be applied directly to this system. The available methodologies to size this system were reviewed. Power available calculation procedure and factors affects it were addressed prior to sizing process. Various design issues of this system due to interrelationship of internal gas flow dynamics and rotor external aerodynamics was discussed. Finally, a modification that is required in existing sizing methodologies was identified and combined approach in sizing process to consider the interrelationship among engine, rotor and blade duct was introduced.

• International Journal of Aeronautical and Space Sciences
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• v.7 no.2
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• pp.86-103
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• 2006

Using a combination of force transducer measurement to quantify net lift force, high frame rate camera to quantify and subtract inertial contributions, and Digital Particle Image Velocimetry (DPIV) to calculate aerodynamic contributions in the spanwise plane, the contribution of spanwise flow to the generation of lift force in wings undergoing a pure flapping motion in hover is shown as a function of flapping angle throughout the flapping cycle. These experiments were repeated at various flapping frequencies and for various wing planform sizes for flat plate and span wise cambered wings. Despite the previous identification of the importance of span wise fluid structures in the generation of lift force in flapping wings throughout the existing body of literature, the direct contribution of spanwise flow to lift force generated has not previously been quantified. Therefore, in the same manner as commonly applied to investigate the chordwise lift distribution across an airfoil in flapping wings, spanwise flow due to bulk flow and rotational fluid dynamic mechanisms will be investigated to validate the existence of a direct component of the lift force originating from the flapping motion in the spanwise plane instead.

• The Journal of Korean Institute of Communications and Information Sciences
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• v.17 no.4
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• pp.361-370
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• 1992

A controller of the hovering VTOL aircraft with four fan is constructed by SRFIMF(State Rate Feedback Implicit Model-Following)theory, in which feedback state are angle acceleration, angle velocity and angle position of the aircraft during hover With yaw control of the system, characteristics of the hovering aircraft can be analyzed by changing states feedback gain and sponse provides robust stable hovering system.

• Proceedings of the KIEE Conference
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• 1999.07b
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• pp.556-559
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• 1999

In this paper, the Observer Controller IDentification/Eigensystem Realization Algorithm(OCID/ERA) is applied to identify the state space model of a helicopter in hover from a given set of the general input-output data. The objective of this paper is the evaluation of the utility of the OCID/ERA approach to the system identification of a system which has natural unstable mode like a helicopter and to give some guidelines before applying the technique to the real system. As it mentioned above, since a model helicopter is unstable naturally, the closed-loop system identification method(OCID/ERA) using excitation, output and feedback signals of the simulated system is applied.

• Proceedings of the KIEE Conference
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• 1999.07b
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• pp.818-821
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• 1999

Helicopter system is regarded as a challenging example in multivariable robust control application since the dynamics of helicopter is highly coupled and nonlinear. In this paper, Dynamic equations for model helicopter at hover are derived. Various system properties are stated with respect to control of the attitude of the vehicle. A linearized model is used to analyse the system stability and to design the attitude controller. The simulation results of LQG controller are presented.