Influence of Leading-Edge Tubercle with Amplitude Modulation on NACA 0015 Airfoil

AIAA Journal ◽  
2021 ◽  
pp. 1-14
Author(s):  
Mehmet Seyhan ◽  
Mustafa Sarioglu ◽  
Yahya E. Akansu
Keyword(s):  
Amplitude Modulation ◽  
Leading Edge ◽  
Naca 0015

Investigation on the effect of leading edge tubercles of sweptback wing at low reynolds number

2020 ◽  
Vol 21 (6) ◽  
pp. 621
Author(s):  
Veerapathiran Thangaraj Gopinathan ◽  
John Bruce Ralphin Rose ◽  
Mohanram Surya
Keyword(s):  
Leading Edge ◽  
Reynolds Numbers ◽  
Sweep Angle ◽  
Laminar Separation ◽  
Low Reynolds Numbers ◽  
Flow Energy ◽  
Airplane Wing ◽  
Low Reynolds ◽  
Naca 0015 ◽  
Wing Performance

Aerodynamic efficiency of an airplane wing can be improved either by increasing its lift generation tendency or by reducing the drag. Recently, Bio-inspired designs have been received greater attention for the geometric modifications of airplane wings. One of the bio-inspired designs contains sinusoidal Humpback Whale (HW) tubercles, i.e., protuberances exist at the wing leading edge (LE). The tubercles have excellent flow control characteristics at low Reynolds numbers. The present work describes about the effect of tubercles on swept back wing performance at various Angle of Attack (AoA). NACA 0015 and NACA 4415 airfoils are used for swept back wing design with sweep angle about 30°. The modified wings (HUMP 0015 A, HUMP 0015 B, HUMP 4415 A, HUMP 4415 B) are designed with two amplitude to wavelength ratios (η) of 0.1 & 0.24 for the performance analysis. It is a novel effort to analyze the tubercle vortices along the span that induce additional flow energy especially, behind the tubercles peak and trough region. Subsequently, Co-efficient of Lift (CL), Co-efficient of Drag (CD) and boundary layer pressure gradients also predicted for modified and baseline (smooth LE) models in the pre & post-stall regimes. It was observed that the tubercles increase the performance of swept back wings by the enhanced CL/CD ratio in the pre-stall AoA region. Interestingly, the flow separation region behind the centerline of tubercles and formation of Laminar Separation Bubbles (LSB) were asymmetric because of the sweep.

scholarly journals Flow Separation Control of Backward-Facing Step Airfoil NACA0015 by Blowing Technique

2019 ◽  
Vol 12 (1) ◽  
pp. 99-119
Author(s):  
Khuder N. Abed
Keyword(s):  
Flow Separation ◽  
Numerical Study ◽  
Leading Edge ◽  
Distribution Coefficients ◽  
Open Circuit ◽  
Experimental Investigations ◽  
Backward Facing Step ◽  
Ansys Fluent ◽  
Flow Separation Control ◽  
Naca 0015

The aim of this paper is to control the flow separation above backward-facing step (BFS) airfoil type NACA 0015 by blowing method. The flow field over airfoil has been studied both experimentally and computationally. The study was divided into two parts: a practical study through which NACA 0015 type with a backward -facing step (located at 44.4% c from leading edge) on the upper surface containing blowing holes parallel to the airfoil chord was used. The tests were done over two-dimensional airfoil in an open circuit suction subsonic wind tunnel with flow velocity 25m/s to obtain the pressure distribution coefficients. A numerical study was done by using ANSYS Fluent software version 16.0 on three models of NACA 0015, the first one has backward-facing step without blowing, the second with single blowing holes and the third have multi blowing holes technique. Both studies (experimental and numerical) were done at low Reynolds number (Re=4.4x105) and all models have chord length 0.27m.The experimental investigations and CFD simulations have been performed on the same geometry dimensions, it has been observed that the flow separation on the airfoil can be delayed by using  velocity blowing (30m/s) on the upper surface. The multi blowing holes with velocity improved the aerodynamics properties.The multi blowing holes and single blowing hole thesame effect onpressure distribution coefficients

scholarly journals Experiment on Wingtip Vortices using a Half Deltawing at the Tips

2019 ◽  
Vol 8 (2S8) ◽  
pp. 1633-1638
Keyword(s):  
Delta Wing ◽  
Tangential Velocity ◽  
Leading Edge ◽  
Force Balance ◽  
Aviation Industry ◽  
Sweep Angle ◽  
Wingtip Vortex ◽  
Image Velocimetry ◽  
Rotating Wing ◽  
Naca 0015

The counter rotating wing tip vortices produced by the aircraft continues to be a big concern for the aviation industry and the aircraft manufacturers due to its hazardous effects on the flight safety and aircraft efficiency. The strength of the vortices poses severe problems to the aircraft operations. Manufacturers developed various wingtip devices to alleviate this problem, but still it is not fully understood and solved. In this thesis, the effectiveness of using a half delta wing at the tips is investigated. The flow field over a low aspect ratio NACA 0015 wing fitted with a slender sharp half delta wing with a leading edge sweep angle 700 at a Reynolds number 1.87 ×105 is investigated. Particle image velocimetry is used to quantify the vortex structure and force balance measurements are used to calculate the aerodynamic data of the wing. The peak vorticity, peak tangential velocity are decreased due to the addition of half delta wing. The over-all radius of the wingtip vortex increased showing a diffused vortex due to the addition of the half delta wing. The core circulation is decreased leading to a lower strength vortex. Though the tip device increased the drag, it increases the aerodynamic efficiency through the improvement in L/D.

scholarly journals Improving the aerodynamic performance of a cycloidal rotor through active compliant morphing

2017 ◽  
Vol 121 (1241) ◽  
pp. 901-915
Author(s):  
L. Ferrier ◽  
M. Vezza ◽  
H. Zare-Behtash
Keyword(s):  
Negative Impact ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Micro Air Vehicles ◽  
Aerodynamic Performance ◽  
Dynamic Stall ◽  
Marine Vehicles ◽  
Aerospace Applications ◽  
Naca 0015 ◽  
Air Vehicles

ABSTRACTCycloidal rotors are a novel form of propulsion system that can be adapted to various forms of transport such as air and marine vehicles, with a geometrical design differing significantly from the conventional screw propeller. Research on cycloidal rotor design began in the early 1930s and has developed throughout the years to the point where such devices now operate as propulsion systems for various aerospace applications such as micro air vehicles, unmanned air vehicles and compound helicopters. The majority of research conducted on the cycloidal rotor’s aerodynamic performance have not assessed mitigating the dynamic stall effect, which can have a negative impact on the rotor performance when the blades operate in the rotor retreating side. A solution has been proposed to mitigate the dynamic stall effect through employment of active, compliant leading-edge morphing. A review of the current state of the art in this area is presented. A two-dimensional, implicit unsteady numerical analysis was conducted using the commercial computational fluid dynamics software package STAR CCM+, on a two-bladed cycloidal rotor. An overset mesh technique, otherwise known as a chimera mesh, was used to apply complex transient motions to the simulations. Active, compliant leading-edge morphing is applied to an oscillating NACA 0015 aerofoil to attempt to mitigate the dynamic stall whilst maintaining the positive dynamic lift coefficient (Cl) contributions. It was verified that by applying a pulsed input leading-edge rotational morphing schedule, the leading-edge vortex does not fully form and the large flow separation is prevented. Further work in this investigation will focus on coupling the active, leading-edge motion to the cycloidal rotor model with the aim to maximise aerodynamic performance.

Vortex Generators Contribution to the Enhancement of the Aerodynamic Performances

2014 ◽  
Vol 950 ◽  
pp. 268-274
Author(s):  
Hocine Tebbiche ◽  
Mohamed S. Boutoudj
Keyword(s):  
Wind Tunnel ◽  
Flow Control ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Vortex Generators ◽  
Drag Coefficients ◽  
Passive Devices ◽  
Aerodynamic Performances ◽  
Lift And Drag ◽  
Naca 0015

This study interest flow control using a new vortex generators (VGs) shape with counter-rotating vortices, obtained by adding a new element to a configuration mostly investigated. The experiments were performed in the aim to determine the VGs answer when placed on the suction face at 10% from the leading edge of an airfoil Naca 0015 in order to improve the lift and drag coefficients. The investigations were accomplished in wind tunnel for two Reynolds numbers and geometrical vortex generators configurations. The obtained results are analyzed according to several parameters such as the VG height, the space between the same VG pair and the additional factor. The results show a profit brought by the passive devices estimated at about 28% of the CL/Cd ratio.

Control of leading edge separation on airfoil using DBD plasma actuator with signal amplitude modulation

2015 ◽  
Vol 19 (1) ◽  
pp. 37-47 ◽  
Author(s):  
Nazri Md Daud ◽  
Yasuaki Kozato ◽  
Satoshi Kikuchi ◽  
Shigeki Imao
Keyword(s):  
Amplitude Modulation ◽  
Leading Edge ◽  
Signal Amplitude ◽  
Plasma Actuator ◽  
Dbd Plasma ◽  
Edge Separation

Blade Offset and Pitch Effects on a High Solidity Vertical Axis Wind Turbine

2009 ◽  
Vol 33 (3) ◽  
pp. 237-246 ◽  
Author(s):  
Andrzej J. Fiedler ◽  
Stephen Tullis
Keyword(s):  
Wind Speed ◽  
Wind Tunnel ◽  
Wind Turbine ◽  
Leading Edge ◽  
Vertical Axis ◽  
Small Scale ◽  
Vertical Axis Wind Turbine ◽  
Data Set ◽  
Naca 0015 ◽  
Base Data

A high solidity, small scale, 2.5m diameter by 3m high Vertical Axis Wind Turbine (VAWT) consisting of three NACA 0015 profile blades, each with a span of 3m and a chord length of 0.4m, was tested in an open-air wind tunnel facility to investigate the effects of preset toe-in and toe-out turbine blade pitch. The effect of blade mount-point offset was also investigated. The results from these tests are presented for a range of tip speed ratios, and compared with an extensive base data set obtained for a nominal wind speed of 10m/s. Results show measured performance decreases of up to 47% for toe-in, and increases of up to 29% for toe-out blade pitch angles, relative to the zero preset pitch case. Also, blade mount-point offset tests indicate decreases in performance as the mount location is moved from mid-chord towards the leading edge, as a result of an inherent toe-in condition. Observations indicate that these performance decreases may be minimized by compensating for the blade mount offset with a toe-out preset pitch. The trends of the preset blade pitch tests agree with those found in literature for much lower solidity turbines.

Closed-Loop Feedback Control Algorithms for Flow Control Over NACA 0015 Airfoil

2011 ◽  
Author(s):  
Sohaib Obeid ◽  
Rataheshwar Jha ◽  
Goodarz Ahmadi
Keyword(s):  
Feedback Control ◽  
Closed Loop ◽  
Moving Average ◽  
Leading Edge ◽  
Open Loop ◽  
Synthetic Jet ◽  
Identification Technique ◽  
Loop Feedback ◽  
Closed Loop Feedback ◽  
Naca 0015

This study investigates control algorithm for closed-loop feedback control system design aimed at reduction of turbulent flow separation over a NACA 0015 airfoil equipped with leading-edge synthetic jet actuators (SJAs). The algorithm employs system identification technique based on Nonlinear Auto Regressive Moving Average with eXogenous inputs (NARMAX) method to model nonlinear dynamics of the flow and design controller for single-input singleoutput systems. The resulting closed loop response tracks the desired pressure value and significant improvement in the transient response over the open-loop system at high angles of attack is realized.

Experimental investigation of flowfield over an iced aerofoil

2016 ◽  
Vol 120 (1227) ◽  
pp. 735-756 ◽  
Author(s):  
M.D. Manshadi ◽  
M.K. Esfeh
Keyword(s):  
Shear Layer ◽  
Surface Pressure ◽  
Low Frequency ◽  
Leading Edge ◽  
Velocity Fluctuations ◽  
Surface Pressure Distribution ◽  
Unsteady Separation ◽  
Flow Features ◽  
Velocity Spectra ◽  
Naca 0015

ABSTRACTWind-tunnel measurements were used to study the characteristics of the unsteady separation bubbles on a NACA 0015 aerofoil with simulated two-dimensional leading-edge glaze ice accretions. The unsteadiness present in the iced-aerofoil flowfield was determined using measurements of the time-dependent aerofoil surface pressure distribution at Reynolds number of 1.0 × 106. Additionally, the unsteady flow features were investigated through the power spectrum of the stream-wise velocity fluctuations using a hot-wire anemometry. The results showed that the highest value of root-mean-square fluctuation of the surface pressure consistently occurred upstream of the mean shear-layer reattachment location. Spectral analysis of stream-wise velocity fluctuations near reattachment location revealed evidence of the regular frequency at Strouhal numbers of 0.5-0.63. Moreover, the low-frequency oscillations associated with shear-layer flapping was also identified in the wake velocity spectra on the order of 10 Hz that resulted in Strouhal numbers of 0.0186-021.

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