Kenichi Rinoie

Experiments on a 60-Degree Delta Wing with Rounded Leading-Edge Vortex Flaps

Low-speed wind-tunnel measurements were done on a 1.15-m span 60-deg delta wing with rounded leading-edge vortex flaps. The purpose of the measurements is to assess the benefits of the rounded leading-edge vortex flaps in regard to improving the lift/drag ratio of delta wings. Force and surface pressure measurements were made at a Reynolds number based on a centerline chord of 2 x 10 6 . The increase in the radius of the rounded leading edge reduces the drag significantly both with and without flap deflection except in the minimum drag region. Deflecting the rounded leading-edge vortex flap improves the lift/drag ratio at relatively higher lift coefficients, when compared with the sharp-edged vortex flap. The largest improvement in the lift/drag ratio as compared with the sharp-edged delta wing with vortex flaps is more than 25% in the lift coefficient range between about 0.6 and 0.8 for the rounded-edge delta wing with flaps that were deflected 30 deg downward

Airfoil Stall Suppression by Use of a Bubble Burst Control Plate

To suppress the stall on an NACA 0012 airfoil, a thin plate (hereafter referred to as a burst control plate) was attached on the airfoil to delay the burst of laminar separation bubbles formed at the stall angle. The burst control plate is used to enhance the vortical structures in the separated shear layer. Wind-tunnel tests were conducted at a chord Reynolds number of 1.3 x 10 5 . Flow visualization tests and surface pressure measurements showed that the burst control plate placed on the airfoil suppresses the bubble burst at a wide range of angle of attack and therefore, both the stall angle and the maximum lift coefficient are increased. The particle image velocimetry measurements indicated that when the angle of attack of the airfoil with the burst control plate is set higher than the stall angle of the original airfoil, a similar flow structure of the original airfoil was observed.

Experimental Studies of a 70-Degree Delta Wing with Vortex Flaps

Force, surface pressure, and flowfield measurements were made on a 0.5-m root chord, 70-deg delta wing model with leading-edge vortex flaps at the National Aerospace Laboratory, Japan. The main objective of the experiment is to investigate the effect of the wing sweepback angle upon the vortex flap performance. Improvements in the lift/drag ratio were observed in a 70-deg delta wing by deflecting the tapered vortex flaps.

Stereoscopic PIV measurements of leading edge separation vortices on a cranked arrow wing

Three-component velocity measurement results via stereoscopic PIV for leeward flow of a cranked arrow wing with a centre body are presented in order to understand leeward flow phenomena, emphasizing angle-of-attack effects and Reynolds number effects. The tests are conducted in a low-speed wind tunnel with the Reynolds number range from 0.15 × 106 to 1.5 × 106 at angles of attack of 8° to 20°. Based on instantaneous results, statistical properties such as turbulence kinetic energy are derived. The test results indicate that with increase in angle of attack, interaction between two leading edge separation vortices and a vortex on the body becomes strong, merging into a single vortex. In results on angle of attack effects, vorticity magnitude and kinetic energy have a positive correlation in the cores of the separation vortices. Results of Reynolds number effects show that with increasing Reynolds number, the secondary vortex structure becomes small with the two primary separation vortices moving outward, and the core diameter and vorticity of the vortices become larger and smaller, respectively, due to boundary layer transition on the inboard wing. In general, these Reynolds number effects are similar to the effects of decrease in angle of attack.

Surface pressure distributions on 4% circular arc airfoil at low Reynolds number

V ERY small aircraft, called micro air vehicles (MAVs), are of high interest because electronic equipment can be miniaturized to allow for the easy manufacture of a vehicle whose entire mass is only a few dozen grams. Small-sized MAVs operate at chord Reynolds numbers below 1:0 10. Because of the influence of viscosity, thin and sharp leading edge airfoils with thickness ratios less than about 5%offer better aerodynamic characteristics than thick and blunt edge airfoils with Reynolds numbers lower than 1:0 10 [1]. Mueller has measured the aerodynamic forces acting on a circular arc airfoil for MAVs; however, compared with thick and blunt edge airfoils, very little literature is available for thin and sharp leading-edge airfoils [1,2], which indicates that more research needs to be done for the low Reynolds number region. Cosyn and Vierendeels [3] numerically studied the low Reynolds number aerodynamics of a flat plate and an S5010 airfoil, which is an airfoil with a 10% thickness ratio. They have also pointed out that low Reynolds number flows exhibit complex flow phenomena, such as laminar separation, which was described by Mueller and DeLurier [4]. For the design and manufacture of MAVs, it is important to know the details of their aerodynamic characteristics, such as surface pressure distribution. However, there is no space inside of the circular arc airfoil for plumbing the pipes from the static pressure port to the pressure transducer. Thus, as far as the authors know, there are no experimental data measuring surface pressure distributions over a circular arc airfoil except for [2], which used luminescent, pressuresensitive paint. In this study, to measure the surface static pressure of the circular arc airfoil, wemade a 4% cambered-airfoil sectionmodel with a 1% thickness ratio by soldering copper pipes. We also conducted surface flow visualizations using the oil flow technique. The present experimental results at a chord Reynolds number of Re 62; 000 are expected to provide useful information for understanding the flowfield of the circular arc airfoil at a low Reynolds number aswell as for confirming the accuracy of numerical estimations concerning a circular arc airfoil.

Studies on Vortex Flaps with Rounded Leading Edges for Supersonic Transport Configuration

Wind-tunnel measurements were taken on a cranked arrow wing supersonic transport configuration with leading-edge vortex flaps. Force and surface pressure measurements were made at Reynolds number based on the wing mean aerodynamic chord from 9.2 × 105 to 3.8 × 10 6 . Two different flap cross sections (the originally designed nonrounded leading edge and the rounded leading edge) were tested. The purpose of the measurements is to clarify how the differences of the Reynolds number affect the flow around the rounded leading-edge vortex flaps and the flap performance. The wing with the rounded leading-edge vortex flaps indicated some benefit of the lift/drag ratio as compared with those of the nonrounded vortex flaps at a relatively high-lift coefficient greater than 0.3

Laminar Airfoil Modification Attaining Optimum Drag Reduction by Use of Airfoil Morphing

. Aerodynamic characteristics of the baseline and deformed airfoils have been investigated using a viscous–inviscid interaction method. It is shown that the leadingedge deformation is effective in reducing the drag at the offdesign angle of attack, in comparison with the baseline airfoil. The transition point has been estimated, using a numerical method based on a linear stability theory. The deformationisaneffectivemeanstomove thetransition pointaftontheairfoil,andtheextension ofthe laminar flow area results in a reduction in the drag at the offdesign angle of attack. Nomenclature Cd = drag coefficient Cl = lift coefficient Cp = pressure coefficient based on the freestream static and dynamic pressures c = airfoil chord length, m l = girth of the airfoil at the leading edge, m n = amplification factor Rx = Reynolds number based on ue and coordinate along the airfoil surface measured from the leading edge R� = Reynolds number based on ue and � Rec = Reynolds number based on the chord length t = airfoil half-thickness distribution, m U1 = freestream velocity, m=s ue = local velocity at the edge of the boundary layer, m=s x = Cartesian coordinate along the chord direction, measured from the leading edge of the baseline airfoil, m y = Cartesian coordinate perpendicular to x and measured from the leading edge, m ycamber = airfoil camber line, m Z = Cartesian coordinate parallel to x and

Bubble Burst Control Using Smart Structure Sensor Actuators for Stall Suppression

Based on our previous experimental work conducted pertinent to the flow separation bubble burst control using a thin plate (hereafter burst control plate) attached to the leading-edge of laminar airfoils, the burst control plate in the current study is incorporated in the smart structure sensor actuators system on a NACA 631-012 airfoil for stall suppression, lift enhancement and drag reduction. The smart structure sensor actuators system uses a pressure transducer, a displacement sensor and a feedback controller to operate and to monitor the actuators in an autonomous mode. Note that the height of the burst control plate depends on the angular movement of the actuators located on either end of the airfoil section. Experimental studies were conducted at two different Reynolds numbers of 1.3 x 10 5 and 1.9 x 10 5 based on the airfoil chord length. The wind tunnel experiments demonstrated the effectiveness of the smart structure sensor actuators system and the burst control plate located at 6% chord-wise position (defined at the trailing-edge of the plate) of the airfoil, in that the drag is reduced, stall angle is delayed, and a significant improvement in the lift-to-drag ratio is achieved over a wide range of angles-of-attack. The feedback control approach provides a means to strengthen the vortical structures inside the separated shear layer with the overall drag of the airfoil is kept minimal. The experimental results demonstrated the application of the smart structure sensor actuators system and the burst control plate on the NACA 631-012 airfoil can be effective means of bubble burst control and airfoil stall suppression in low speed air vehicle applications.

Three-Dimensional Separated Flow on a Flat Plate with Leading-Edge Serrations

Serrations, which imitate the sawteeth-like structures in the leading edge of owl wing, enable to reduce noises and alter the aerodynamic performance. However, the mechanism is not yet completely understood because the flow fields have three-dimensionality. In this study, in order to investigate three-dimensional mean flow structures on a flat plate with leadingedge serrations, flow fields viewed from the spanwise and vertical directions were measured by two-dimensional time-resolved PIV and stereo PIV at the Reynolds number of 5.0×10. When the angle of attack was 6 degree, the interval of the flow pattern was equal to the width of one sawtooth of the serrations. Whereas the flow over the trough was separated, and the flow over the vertex passed along the surface of the flat plate. This is because pairs of longitudinal vortices induced by the serrations suppress the flow over the vertex on the uppersurface of the flat plate. When the angle of attack was 14 degree, several flows over the vertexes were gathered, and specific flow patterns were formed. The flow over the vertex passed along the surface of the flat plate, so that separation points viewed from the spanwise direction moved to the downstream side. Reverse-flow regions behind the troughs could be divided into two categories, “open” and “closed” ones. Pairs of longitudinal vortices were seen in the closed reverse-flow regions only, and growth of the open reverse-flow regions forces the vortices to move upwards.

Bubble Burst Control for Stall Suppression on a NACA 631-012 Airfoil

Based on the novel concepts of the burst control plate, the current article aims to demonstrate the effectiveness of the burst control plate attached at different locations (5%, 7.5% and 10% of the chord of the airfoil) on a NACA 631-012 airfoil section for laminar separation bubble burst delay, stall suppression, lift augmentation and drag reduction at a chord Reynolds number of 1.3x10 5 . The effectiveness of the burst control plate is governed by its height, width, geometry and the distance between the leading-edge of the airfoil and the trailing-edge of the plate. Flow patterns around the airfoil with and without the plate attachment were studied with an aid of smoke flow visualization. The results suggest that the stall angle of the original airfoil (without the burst control plate) occurs at 10 o and it is successfully postponed to approximately 13 o when the burst control plate is attached onto the airfoil. Lift and drag forces were measured for a range of angle-of-attack from 0 o to +20 o . The experimental results have shown that the lift generated on the NACA 631-012 airfoil with the plate attachment and above 9 o angle-of-attack is significantly higher than the original airfoil, while the drag is sufficiently reduced. The overall results demonstrated the application of the burst control plate on the NACA 631-012 airfoil can be effective means of bubble burst control and airfoil stall suppression in low speed flows.