• International Journal of Aeronautical and Space Sciences
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• v.13 no.3
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• pp.317-322
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• 2012

The low speed aerodynamic characteristics for a modified sonic arc airfoil which is designed by using the nose shape function of sonic arc, the shape function of NACA four-digit wing sections, and Maple are experimentally investigated. The small rotor blades of a modified sonic arc and NACA0012 airfoil are precisely fabricated with a commercially available light aluminum(Al 6061-T6) and are spin tested over a low speed range (3000rpm-5000rpm). In a consuming power comparison, the consuming powers of NACA0012 are higher than that of modified sonic arcs at each pitch angle. The measured rotor thrust for each pitch angle is used to estimate the rotor thrust coefficient according to momentum theory in the hover state. The value of thrust coefficients for both two airfoils at each pitch angle show almost constant values over the low Mach number range. However, the rotor thrust coefficient of NACA0012 is higher than that of the modified sonic arc at each pitch angle. In conclusion, the aerodynamic performance of NACA0012 is better than that of modified sonic arcs in the low speed regime. This test model will provide a convenient platform for improving the aerodynamic performance of small scale airfoils and for performing design optimization studies.

• International Journal of Aeronautical and Space Sciences
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• v.18 no.4
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• pp.601-613
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• 2017

An adjoint-based error estimation and mesh adaptation study is conducted for two-dimensional viscous flows on unstructured hybrid meshes. The error in an integral output functional of interest is estimated by a dot product of the residual vector and adjoint variable vector. Regions for the mesh to be adapted are selected based on the amount of local error at each nodal point. Triangular cells in the adaptive regions are refined by regular refinement, and quadrangular cells near viscous walls are bisected accordingly. The present procedure is applied to single-element airfoils such as the RAE2822 at a transonic regime and a diamond-shaped airfoil at a supersonic regime. Then the 30P30N multi-element airfoil at a low subsonic regime with a high incidence angle (${\alpha}=21deg.$) is analyzed. The same level of prediction accuracy for lift and drag is achieved with much less mesh points than the uniform mesh refinement approach. The detailed procedure of the adjoint-based mesh refinement for the multi-element airfoil case show that the basic flow features around the airfoil should be resolved so that the adjoint method can accurately estimate an output error.

• Aerospace Engineering and Technology
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• v.3 no.2
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• pp.248-255
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• 2004

In this study, aerodynamic shape design of airfoils was performed by using RSM(response surface method) and two-dimensional Navier-Stokes solver. Numerical experiment points were determined by D-optimal method and quadratic response surfaces were constructed by using JMP. For the validations of design method, NACA 64621 airfoil was inversely designed to have aerodynamic characteristics of Bell airfoil. The design method was applied to the aerodynamic design of both smart UAV wing airfoil and low Reynolds rotor-blade airfoil for unmanned helicopter. The optimized airfoils showed improved performance with various constraint conditions.

• 한국전산유체공학회:학술대회논문집
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• 2006.10a
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• pp.193-198
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• 2006

It takes a lot of time and effort to generate grids for numerical analysis of problems with ground effect because the relative attitude and height of airfoil should be maintained to the ground as well as the inflow. A low Mach number preconditioned turbulent flow solver using the overlap grid technique has been developed and applied to the ground effect simulation. It has been validated that the present method using the multi-block grid gives us highly accurate solutions comparing with the experimental data of the RAE 101 airfoil in an unbounded condition. Present numerical method has been extended to simulate ground effect problems by using the overlapped grid system to avoid tedious work in generating multi-block grid system. An extended method using the overlapped grid has been verified and validated by comparing with results of multi-block method and experimental data as well. Consequently, the overlapped grid method can provide not only sufficiently accurate solutions but also the efficiency to simulate ground effect problems. It is shown that the pressure and aerodynamic centers move backward by the ground effect as the airfoil approaches to the ground.

• Proceedings of the KSME Conference
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• 2002.08a
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• pp.479-482
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• 2002

Riblets with longitudinal grooves along the streamwise direction have been used as an effective flow control technique for drag reduction. A flexible micro-riblet with v-grooves of peak-to-peak spacing of $300{\mu}m$ was made using a MEMS fabrication process of PDMS replica. The flexible micro-riblet was attached on the whole surface of a NACA0012 airfoil with which grooves are aligned with the streamwise direction. The riblet surface reduces drag coefficient about $7.9{\%}\;at\;U_o=3.3m/s$, however, it increases drag about $8{\%}\;at\;U_o=7.0m/s$, compared with the smooth airfoil without riblets. The near wake has been investigated experimentally far the cases of drag reduction ($U_o\;=\;3.3 m/s$) and drag increase ($U_o\;=\;7 m/s$). Five hundred instantaneous velocity fields were measured for each experimental condition using the cross-correlation PIV velocity field measurement technique. The instantaneous velocity fields were ensemble averaged to get spatial distribution of turbulent statistics such as turbulent kinetic energy. The experimental results were compared with those of a smooth airfoil under the same flow condition. The micro-riblet surface influences the near wake flow structure largely, especially in the region near the body surface

• Transactions of the Korean Society of Mechanical Engineers B
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• v.26 no.5
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• pp.685-694
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• 2002

Effects of wake-passing orientation and frequency on the wake-induced boundary layer transition on a NACA0012 airfoil are investigated. The wakes are generated by rotating cylinders clockwise (CW) and counterclockwise (CCW) around the airfoil. Time- and phase-averaged streamwise mean velocities and turbulent fluctuations are measured with a single hot-wire probe. Wall skin frictions are estimated by the Computational Preston Tube Method (CPM). The pressure distribution on the airfoil is different according to the wake-passing orientation and frequency. Turbulent patches are generated in the laminar boundary layer due to the passing wake and the boundary layer becomes temporarily transitional. The transition process is significantly affected by the pressure gradient and the turbulent patches. For the receding wake, the turbulent patches propagate more rapidly than those for the approaching wake because adverse pressure gradient becomes larger. As the frequency increases, onset location of transition moles upstream and the boundary layer near the trailing edge becomes more transitional.

• Advances in aircraft and spacecraft science
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• v.2 no.4
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• pp.427-443
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• 2015

The present paper is the follow-on of a former work in which the influence of the gas-surface interaction models was evaluated on the aerodynamic coefficients of an aero-space-plane and on a section of its wing. The models by Maxwell and by Cercignani-Lampis-Lord were compared by means of Direct Simulation Monte Carlo (DSMC) codes. In that paper the diffusive, fully accommodated, semi-specular and specular accommodation coefficients were considered. The results pointed out that the influence of the interaction models, considering the above mentioned accommodation coefficients, is pretty strong while the Cercignani-Lampis-Lord and the Maxwell models are practically equivalent. In the present paper, the comparison of the same models is carried out considering the dependence of the accommodation coefficients on the angle of incidence (or partial accommodation coefficients). More specifically, the normal and the tangential momentum partial accommodation coefficients, obtained experimentally by Knetchel and Pitts, have been implemented. Computer tests on a NACA-0012 airfoil have been carried out by the DSMC code DS2V-64 bits. The airfoil, of 2 m chord, has been tested both in clean and flapped configurations. The simulated conditions were those at an altitude of 100 km where the airfoil is in transitional regime. The results confirmed that the two interaction models are practically equivalent and verified that the use of the Knetchel and Pitts coefficients involves results very close to those computed considering a diffusive, fully accommodated interaction both in clean and flapped configurations.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.27 no.3
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• pp.334-346
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• 2003

An experimental study was carried out to investigate near-wake characteristics of an elliptic airfoil oscillating in pitch. The airfoil was sinusoidally pitched about the half chord point between -5$^{\circ}$and +25$^{\circ}$angles of attack at the freestream velocities of 3.4 and 23.1 m/s. The corresponding Reynolds numbers based on the chord length were 3.3$\times$10$_{4}$ and 2.2$\times$10$^{5}$ , respectively. A hot-wire anemometer was used to measure the near-wake flow variables at the reduced frequency of 0.1. Ensemble-averaged velocity and turbulence intensity profiles were presented to examine the near-wake characteristics depending on the Reynolds number. The axial velocity deficit in the near-wake region tends to decrease with the increase in the Reynolds number as found in many stationary airfoil tests. Turbulence intensity in the near-wake region have a tendency to decrease with the -increase in the Reynolds number during the pitch-up motion, whereas it shows different feature during the pitch-down motion according to the separation characteristics.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.24 no.8
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• pp.1112-1118
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• 2000

The objective of this study is to increase lift and decrease drag of an airfoil at high angles of attack by delaying flow separation with piezo-ceramic actuators. The airfoil used is NACA 0012 and its chord length is 0.3m. An experiment is performed at the freestream velocity of 15m/s at which the Reynolds number based on the chord length is $2{\times}10^5$. Seven rectangular actuators are attached to the airfoil surface and move up and down based on the electric signal. Drag and lift are measured using an in-house two-dimensional force-balance and the surface pressures are also measured. At the attack angle of $16^{\circ}$, the separation point is delayed downstream due to momentum addition induced by the movement of the actuators. Lift is increased by 10%, drag is reduced by 37%, and the efficiency is increased up to 170%. The flow fields with and without control are visualized using the smoke-wire and tuft techniques.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.25 no.5
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• pp.645-652
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• 2001

Heat transfer data are presented which describe characteristics of the transitional thermal boundary layers on the NACA0012 airfoil with upstream wakes. The wakes are generated periodically by circular cylindrical rods which rotate around the airfoil like a squirrel cage. The unsteady wakes simulate those produced by the upstream rotating blade rows in axial turbomachines. The pressure or suction side of the airfoil is also simulated according to the rotating direction of circular rods. As the Reynolds number and the number of rotating rods increase, the boundary layer transition occurs earlier and the Nusselt number increases. The difference of heat transfer coefficient is less on the pressure side than on the suction side. At a constant Reynolds number, the Nusselt number is larger and smaller, respectively, before and after transition as the Strouhal number increases.