Dong-Ha Kim

Boundary Layer and Near-Wake Measurements of NACA 0012 Airfoil at Low Reynolds Numbers

A study on the boundary layers and near-wake behavior of a NACA 0012 airfoil at low Reynolds numbers was conducted in order to gain knowledge about flow properties applicable to turbine blades and MAVs. Hot-wire anemometers were used to measure the boundary layers and near-wakes at angles-of-attack of α =0°, 3°, and 6°; and Reynolds Numbers of Rec=2.3×10, 3.3×10, and 4.8×10. The boundary layers were measured with small intervals along the direction normal to the airfoil surface by a 3-D traverse system. The near-wakes were measured at four downstream stations of x/C=0.1, 0.3, 0.5, and 1.0. The results show that the laminar separation of the boundary layers occurred at α =3°. The reattachment of the separated boundary layers was clearly observed for the case of Re=4.8×10 at α =6°, showing the formation of a long separation bubble. The properties of the long separation bubble were then observed and found to be different compared to the properties of short separation bubbles studied previously. The shear flow instability wave was investigated with respect to the formation of coherent vortical structures in the separated shear layer. It was found that the Reynolds number critically influenced the shear layer instability wave rather than the angle-of-attack. The near-wakes in the present paper were analyzed in detail from the boundary layer characteristics that directly affect nearwakes. In particular, the near-wake properties for Rec=4.8×10 and α =6° represented a clearly different evolutions compared to the other Reynolds number cases owing to reattachment.

Reynolds Number Effects on Unsteady Boundary Layer of an Oscillating Airfoil

An experimental study was conducted in order to investigate the influence of Reynolds number on an oscillating airfoil. A NACA 0012 airfoil was sinusoidally pitched at quarter chord, and the oscillation amplitude was between ±6°. Surface-mounted probes and smokewire visualization were used for the examination of the boundary layers. The free-stream velocities were 1.98, 2.83, and 4.03 m/s, and the corresponding chord Reynolds numbers were Rec=2.3×10, 3.3×10, and 4.8×10, respectively. In each case, the frequency of the airfoil oscillation was adjusted to fix the reduced frequency of 0.1. It was observed in the presented low Reynolds number range that for a pitching airfoil, the separation bubble is formed within the stall angle as reattachment occurs. Also, unsteady laminar separation was observed at a saddle point, which is composed by the wall, external flow, and two vortices that is independent of reverse flow. It stands for that laminar separation, reattachment, and unsteady laminar separation occur simultaneously over the airfoil, which means that the unsteady laminar separation occurred in the process of transition after the reattachment of separated boundary layer in unsteady flow. The reattachment and the unsteady laminar separation are promoted (delayed) by the increase of Reynolds number in upstroke (downstroke).

Aerodynamic Characteristics of a Pitching Airfoil Through Pressure-Distortion Correction in Pneumatic Tubing

The lift and pressure-drag coefficients integrated from unsteady-pressure coefficients are investigated for a pitching airfoil at a low Reynolds number. The unsteady pressure on the pitching-airfoil surface is measured through pressure-distortion correction in the pneumatic tubing, and the pressure-distortion characteristics are estimated. The unsteady pressure on the pitching airfoil is significantly delayed with reduced frequency, and the nondimensional time delay is evaluated with respect to unsteady pressure. The hysteresis loops of the lift and pressure-drag coefficients rotate in counterclockwise and clockwise directions, respectively, and this is due to the reattachment during downstroke. Moreover, the angle of attack range, in which aerodynamic performance is nonlinearly varied, is generated due to the nonlinear boundary-layer behavior caused by the low-Reynolds-number effect. As a result, the unsteady aerodynamic characteristics at Rec=4.8×104 are dominantly influenced by the reattachment and the...

Flow disturbances depending on excitation frequency on a flat plate by a piezoceramic actuator

Flow disturbances by a piezoceramic actuator were investigated to obtain physical insights on the streamwise vortex formation and excitation frequency effects on a piezoceramic actuator. Boundary-layer measurements in wind tunnel and flow visualization in water tunnel were carried out to explore flow disturbance. Findings revealed that streamwise vortex pair was formed in the moderate excitation frequency range, accompanied by low- and high-speed regions formation in the streamwise mean velocity field. U-turbulence intensity was tremendously increased in the high excitation frequency range. As a result, the formation of streamwise vortex pair was significantly influenced by excitation frequency and thus, the flow control purpose should be considered when constructing the proper design parameters for a piezoceramic actuator.Graphical Abstract

Influence of Flow Conditions on a Boundary layer to the Near-Wake of a Flat Plate

An experimental study was carried out to investigate the influence of flow conditions on a boundary layer to the near-wake of a flat plate. The flow conditions in the vicinity of trailing edge that is influenced by upstream condition history are an essential factor that determines the physical characteristics of a near-wake. Various tripping wires were used to change flow conditions on a boundary layer near trailing edge. Measurement were made at a freestream velocity of 6.0 m/sec. I-type (55P14) and X-type hot-wire probe (55P61) were used to measure the boundary layer and near-wake, respectively. Both velocity profiles of the boundary layer normalized by friction velocity and hot-wire signal show flow types such as laminar, transition, and turbulent boundary layer at 0.95C from the leading edge. The symmetric wake collapses well to universal equation for center of wake except outer part of wake. The asymmetric wake characteristic depends on initial conditions of trailing edge and it has similar parameters on mean velocity itself in our measured range. However the symmetric universal equation is not useful over asymmetric wake in our measured range.

Drag Coefficient Variations of an Oscillating NACA 0012 Airfoil

An experimental study was performed in order to investigate the influence of Reynolds number on the drag coefficient variations of an oscillating airfoil. A NACA 0012 airfoil was sinusoidally pitched at the quarter chord point with an oscillating amplitude of ±6°. The free-stream velocities were 1.98, 2.83 and 4.03 m/s and the corresponding chord Reynolds numbers were 2.3×104, 3.3×104 and 4.8×104, respectively. The drag coefficient was calculated from the ensemble average velocity measured by an X-type hot-wire probe(X-type, 55R51) in the near-wakes region. In the case of Re

Near-wake Measurements of an Oscillating NACA 0012 Airfoil

An experimental study was carried out in order to investigate the influence of Reynolds number on the near-wake of an oscillating airfoil. An NACA 0012 airfoil was sinusoidally pitched at the quarter chord point, and is oscillated over a range of instantaneous angles of attack of 6. An X-type hot-wire probe was employed to measure the near-wake of an oscillating airfoil, and the smoke-wire visualization technique was used to examine the flow properties of the boundary layer. The free-stream velocities were 1.98, 2.83 and 4.03 m/s and the corresponding chord Reynolds numbers were 2.3104, 3.3104 and 4.8104, respectively. The frequency of airfoil oscillation was adjusted to fix a reduced frequency of K

An Experimental Study of Mutual Relation between Wake and Boundary Layer of a Flat Plate; Mean Velocity Field

An experimental study was carried out to investigate the influence of flow conditions of a boundary layer on the near-wake of a flat plate. Various attaching positions of tripping wires were selected to change flow conditions on a boundary layer. Laminar, transitional, and turbulent boundary layer conditions at 0.98C from the leading edge are imposed to investigate the evolution of symmetric and asymmetric wake. An x-type hot-wire probe(55P61) is employed to measure at 8 stations of the near-wake region. Measured mean velocity distributions are presented in terms of similarity parameter. It is found that the symmetric wake collapses well to the universal profile in the central part of the wake. However, the universal profile is not suitable in describing an asymmetric wake.