MEASUREMENTS OF DEFORMATION OF THE ADAPTIVE WING LEADING EDGE IN A WIND TUNNEL BY THE VIDEOGRAMMETRY METHOD

Experimental investigations of wind turbine blades having NACA airfoils 0021 and 4412 with and without tubercles on the leading edge have been performed in a wind tunnel. It was found that the lift coefficient of the airfoil 0021 with tubercles was higher at Re = 1.2×105 and 1.69×105 in post critical region (at higher angle of attach) than airfoils without tubercles but this difference relatively diminished at higher Reynolds numbers and beyond indicating that there is no effect on the lift coefficients of airfoils with tubercles at higher Reynolds numbers whereas drag coefficient remains unchanged. It is noted that at Re = 1.69×105, the lift coefficient of airfoil without tubercles drops from 0.96 to 0.42 as the angle of attack increases from 15° to 20° which is about 56% and the corresponding values of lift coefficient for airfoil with tubercles are 0.86 and 0.7 at respective angles with18% drop.

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Comparison of Wind Tunnel Airfoil Performance Data With Wind Turbine Blade Data

Journal of Solar Energy Engineering ◽

10.1115/1.2929989 ◽

1992 ◽

Vol 114 (2) ◽

pp. 119-124 ◽

Cited By ~ 12

Author(s):

C. P. Butterfield ◽

George Scott ◽

Walt Musial

Keyword(s):

Wind Tunnel ◽

Wind Turbine ◽

Peak Power ◽

Unsteady Aerodynamics ◽

Leading Edge ◽

Operating Conditions ◽

Performance Data ◽

Horizontal Axis Wind Turbine ◽

Force Coefficients ◽

Basic Fluid

Horizontal axis wind turbine (HAWT) performance is usually predicted by using wind tunnel airfoil performance data in a blade element momentum theory analysis. This analysis assumes that the rotating blade airfoils will perform as they do in the wind tunnel. However, when stall-regulated HAWT performance is measured in full-scale operation, it is common to find that peak power levels are significantly greater than those predicted. Pitch-controlled rotors experience predictable peak power levels because they do not rely on stall to regulate peak power. This has led to empirical corrections to the stall predictions. Viterna and Corrigan (1981) proposed the most popular version of this correction. But very little insight has been gained into the basic cause of this discrepancy. The National Renewable Energy Laboratory (NREL), funded by the DOE, has conducted the first phase of an experiment which is focused on understanding the basic fluid mechanics of HAWT aerodynamics. Results to date have shown that unsteady aerodynamics exist during all operating conditions and dynamic stall can exist for high yaw angle operation. Stall hysteresis occurs for even small yaw angles and delayed stall is a very persistent reality in all operating conditions. Delayed stall is indicated by a leading edge suction peak which remains attached through angles of attack (AOA) up to 30 degrees. Wind tunnel results show this peak separating from the leading edge at 18 deg AOA. The effect of this anomaly is to raise normal force coefficients and tangent force coefficients for high AOA. Increased tangent forces will directly affect HAWT performance in high wind speed operation. This report describes pressure distribution data resulting from both wind tunnel and HAWT tests. A method of bins is used to average the HAWT data which is compared to the wind tunnel data. The analysis technique and the test set-up for each test are described.

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Accurate Measurements of Wall Shear Stress on a Plate with Elliptic Leading Edge

Sensors ◽

10.3390/s18082682 ◽

2018 ◽

Vol 18 (8) ◽

pp. 2682 ◽

Cited By ~ 2

Author(s):

Guang-Hui Ding ◽

Bing-He Ma ◽

Jin-Jun Deng ◽

Wei-Zheng Yuan ◽

Kang Liu

Keyword(s):

Shear Stress ◽

Wind Tunnel ◽

Wall Shear Stress ◽

Leading Edge ◽

Reynold Number ◽

Micro Electro Mechanical System ◽

Flow Speed ◽

Silicon On Insulator ◽

Micro Sensor ◽

Wall Shear

A micro-floating element wall shear stress sensor with backside connections has been developed for accurate measurements of wall shear stress under the turbulent boundary layer. The micro-sensor was designed and fabricated on a 10.16 cm SOI (Silicon on Insulator) wafer by MEMS (Micro-Electro-Mechanical System) processing technology. Then, it was calibrated by a wind tunnel setup over a range of 0 Pa to 65 Pa. The measurements of wall shear stress on a smooth plate were carried out in a 0.6 m × 0.6 m transonic wind tunnel. Flow speed ranges from 0.4 Ma to 0.8 Ma, with a corresponding Reynold number of 1.05 × 106~1.55 × 106 at the micro-sensor location. Wall shear stress measured by the micro-sensor has a range of about 34 Pa to 93 Pa, which is consistent with theoretical values. For comparisons, a Preston tube was also used to measure wall shear stress at the same time. The results show that wall shear stress obtained by three methods (the micro-sensor, a Preston tube, and theoretical results) are well agreed with each other.

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Study on the Relation Between Side Ratios of Rectangular Cross Sections and Secondary Vortices at Trailing Edge in Motion-Induced Vortex Excitation

Volume 4: Fluid-Structure Interaction ◽

10.1115/pvp2017-65565 ◽

2017 ◽

Cited By ~ 1

Author(s):

Kazutoshi Matsuda ◽

Kusuo Kato ◽

Kouki Arise ◽

Hajime Ishii

Keyword(s):

Wind Speed ◽

Wind Tunnel ◽

Cross Sections ◽

Leading Edge ◽

Trailing Edge ◽

Wind Speeds ◽

Smoke Flow ◽

Flow Visualizations ◽

Institute Of Technology ◽

Secondary Vortices

According to the results of conventional wind tunnel tests on rectangular cross sections with side ratios of B/D = 2–8 (B: along-wind length (m), D: cross-wind length (m)), motion-induced vortex excitation was confirmed. The generation of motion-induced vortex excitation is considered to be caused by the unification of separated vortices from the leading edge and secondary vortices at the trailing edge [1]. Spring-supported test for B/D = 1.18 was conducted in a closed circuit wind tunnel (cross section: 1.8 m high×0.9 m wide) at Kyushu Institute of Technology. Vibrations were confirmed in the neighborhoods of reduced wind speeds Vr = V/fD = 2 and Vr = 8 (V: wind speed (m/s), f: natural frequency (Hz)). Because the reduced wind speed in motion-induced vortex excitation is calculated as Vr = 1.67×B/D = 1.67×1.18 = 2.0 [1], vibrations around Vr = 2 were considered to be motion-induced vortex excitation. According to the smoke flow visualization result for B/D = 1.18 which was carried out by the authors, no secondary vortices at the trailing edge were formed, although separated vortices from the leading edge were formed at the time of oscillation at the onset wind speed of motion-induced vortex excitation, where aerodynamic vibrations considered to be motion-induced vortex excitation were confirmed. It was suggested that motion-induced vortex excitation might possibly occur in the range of low wind speeds, even in the case of side ratios where secondary vortices at trailing edge were not confirmed. In this study, smoke flow visualizations were performed for ratios of B/D = 0.5–2.0 in order to find out the relation between side ratios of rectangular cross sections and secondary vortices at trailing edge in motion-induced vortex excitation. The smoke flow visualizations around the model during oscillating condition were conducted in a small-sized wind tunnel at Kyushu Institute of Technology. Experimental Reynolds number was Re = VD/v = 1.6×103. For the forced-oscillating amplitude η, the non-dimensional double amplitudes were set as 2η/D = 0.02–0.15. Spring-supported tests were also carried out in order to obtain the response characteristics of the models.

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Mechanically Amplified Piezoelectric Microactuators for Laminar-Turbulent Transition Control on Airfoils

Journal of Microelectronics and Electronic Packaging ◽

10.4071/1551-4897-6.4.211 ◽

2009 ◽

Vol 6 (4) ◽

pp. 211-218 ◽

Cited By ~ 1

Author(s):

C. Bolzmacher ◽

X. Riedl ◽

J. Leuckert ◽

M. Engert ◽

K. Bauer ◽

...

Keyword(s):

Boundary Layer ◽

Wind Tunnel ◽

High Speed ◽

Leading Edge ◽

Transition Process ◽

Open Loop ◽

Film Sensor ◽

Transition Control ◽

Transitional Boundary Layer ◽

Tollmien Schlichting

Drag reduction on airfoils using arrays of micro-actuators is one application of so-called Aero-MEMS. These microactuators interact with TS instabilities (Tollmien-Schlichting waves) inside a transitional boundary layer by superimposing artificially generated counterwaves in order to delay the transition process. These actuators need to exhibit a relatively large stroke at relatively high operational frequencies when operated at high Mach numbers. For this purpose, a novel micromachined mechanical amplification unit for increasing the stroke of piezoelectric microactuators up to high frequencies is proposed. The mechanical lever is provided by a sliced nickel titanium membrane. In this work, the actuator is explained in detail and wind tunnel experiments are carried out to investigate the effect of this mechanically amplified piezoelectric microactuator on thin transitional boundary layers. The experiments have been carried out in the transonic wind tunnel facility of the Berlin University of Technology on an unswept test wing with an NACA 0004 leading edge. The effectiveness of the actuator for flow control applications is determined in an open-loop setup consisting of one actuator having a relevant spanwise extension and a microstructured hot film sensor array located downstream. The aerodynamic results at Mach 0.33 are presented and discussed. It is shown that the actuator influences TS wave specific frequencies between 2.5 kHz and 7.4 kHz. The actuator amplitude is large enough to influence a transitional boundary layer significantly without bypassing the natural transition process which makes this type of micromachined actuator a candidate for high speed TS-control.

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Aeroelastic Model Development for the Integrated Adaptive Wing Technology Maturation Project Wind-Tunnel Test

AIAA AVIATION 2020 FORUM ◽

10.2514/6.2020-2717 ◽

2020 ◽

Author(s):

Josiah Waite ◽

Robert E. Bartels ◽

Bret Stanford

Keyword(s):

Wind Tunnel ◽

Model Development ◽

Wind Tunnel Test ◽

Aeroelastic Model ◽

Adaptive Wing

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Exploration on application of dredging thermal protection in the leading edge of the wing

International Journal of Modern Physics B ◽

10.1142/s0217979220401050 ◽

2020 ◽

Vol 34 (14n16) ◽

pp. 2040105

Author(s):

Xiao-Jun Zhu ◽

Feng Li

Keyword(s):

High Temperature ◽

Wind Tunnel ◽

Heat Pipe ◽

Thermal Protection ◽

Wind Tunnel Test ◽

Leading Edge ◽

Temperature Zone ◽

Temperature Heat ◽

Thermal Protection Structure ◽

High Temperature Heat Pipe

Aiming at the severe aerodynamic heating problem in the leading edge of the hypersonic vehicle, in order to ensure the sharp shape of the leading edge of the wing, a dredging thermal protection structure is proposed, and the built-in high-temperature heat pipe structure is used to provide thermal protection for the leading edge of the wing. By means of numerical simulation and arc wind tunnel test, the dredging thermal protection structure of the leading edge of the wing is analyzed, and the thermal protection effect of the built-in high-temperature heat pipe is obtained. The numerical results show that under certain thermal conditions, the temperature at the leading edge of the wing decreases by 304 K, and the minimum temperature of the tail increases by 130 K. The heat flow is dredged from the high-temperature zone to the low-temperature zone, and the thermal load of the leading edge of the wing is weakened. The same result can be obtained by the arc wind tunnel test, which verifies the accuracy of the numerical method and the feasibility of the dredging thermal protection structure with high-temperature heat pipe embedded in the leading edge of the wing.

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Design and Validation of a New Morphing Camber System by Testing in the Price—Païdoussis Subsonic Wind Tunnel

Aerospace ◽

10.3390/aerospace7030023 ◽

2020 ◽

Vol 7 (3) ◽

pp. 23 ◽

Cited By ~ 1

Author(s):

David Communier ◽

Ruxandra Mihaela Botez ◽

Tony Wong

Keyword(s):

Wind Tunnel ◽

Unmanned Aerial Vehicle ◽

Lift Coefficient ◽

Angle Of Attack ◽

Leading Edge ◽

Trailing Edge ◽

Subsonic Wind Tunnel ◽

Aerial Vehicle ◽

Stall Angle

This paper presents the design and wind tunnel testing of a morphing camber system and an estimation of performances on an unmanned aerial vehicle. The morphing camber system is a combination of two subsystems: the morphing trailing edge and the morphing leading edge. Results of the present study show that the aerodynamics effects of the two subsystems are combined, without interfering with each other on the wing. The morphing camber system acts only on the lift coefficient at a 0° angle of attack when morphing the trailing edge, and only on the stall angle when morphing the leading edge. The behavior of the aerodynamics performances from the MTE and the MLE should allow individual control of the morphing camber trailing and leading edges. The estimation of the performances of the morphing camber on an unmanned aerial vehicle indicates that the morphing of the camber allows a drag reduction. This result is due to the smaller angle of attack needed for an unmanned aerial vehicle equipped with the morphing camber system than an unmanned aerial vehicle equipped with classical aileron. In the case study, the morphing camber system was found to allow a reduction of the drag when the lift coefficient was higher than 0.48.

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Airfoil Selection Procedure, Wind Tunnel Experimentation and Implementation of 6DOF Modeling on a Flying Wing Micro Aerial Vehicle

Micromachines ◽

10.3390/mi11060553 ◽

2020 ◽

Vol 11 (6) ◽

pp. 553 ◽

Cited By ~ 1

Author(s):

Taimur Ali Shams ◽

Syed Irtiza Ali Shah ◽

Ali Javed ◽

Syed Hossein Raza Hamdani

Keyword(s):

Wind Tunnel ◽

Equations Of Motion ◽

Selection Procedure ◽

Leading Edge ◽

Control Surface ◽

Flight Tests ◽

Micro Aerial Vehicle ◽

Aerial Vehicle ◽

P Factor ◽

Control Surfaces

Airfoil selection procedure, wind tunnel testing and an implementation of 6-DOF model on flying wing micro aerial vehicle (FWMAV) has been proposed in this research. The selection procedure of airfoil has been developed by considering parameters related to aerodynamic efficiency and flight stability. Airfoil aerodynamic parameters have been calculated using a potential flow solver for ten candidate airfoils. Eppler-387 proved to be the most efficient reflexed airfoil and therefore was selected for fabrication and further flight testing of vehicle. Elevon control surfaces have been designed and evaluated for longitudinal and lateral control. The vehicle was fabricated using hot wire machine with EPP styrofoam of density 50 Kg/ m 3 . Static aerodynamic coefficients were evaluated using wind tunnel tests conducted at cruise velocity of 20 m/s for varying angles of attack. Rate derivatives and elevon control derivatives have also been calculated. Equations of motion for FWMAV have been written in a body axis system yielding a 6-DOF model. It was found during flight tests that vehicle conducted coordinated turns with no appreciable adverse yaw. Since FWMAV was not designed with a vertical stabilizer and rudder control surface, directional stability was therefore augmented through winglets and high wing leading edge sweep. Major problems encountered during flight tests were related to left rolling tendency. The left roll tendency was found inherent to clockwise rotating propeller as ‘P’ factor, gyroscopic precession, torque effect and spiraling slipstream. To achieve successful flights, many actions were required including removal of excessive play from elevon control rods, active actuation of control surfaces, enhanced launch speed during take off, and increased throttle control during initial phase of flight. FWMAV flew many successful stable flights in which intended mission profile was accomplished, thereby validating the proposed airfoil selection procedure, modeling technique and proposed design.

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