The Effect of Pitch Location on Dynamic Stall

1989 ◽  
Vol 111 (3) ◽  
pp. 256-262 ◽  
Author(s):  
E. J. Jumper ◽  
R. L. Dimmick ◽  
A. J. S. Allaire
Keyword(s):  
Wind Tunnel ◽  
Angle Of Attack ◽  
Dynamic Stall ◽  
Stream Velocity ◽  
Pressure Drag ◽  
Constant Rate ◽  
Theoretical Predictions ◽  
Momentum Integral ◽  
Dimensional Variable ◽  
Naca 0015

This paper reports the results of theoretical and wind-tunnel studies of the effect of pitch location on dynamic stall for an airfoil pitching at constant rate. A modified momentum-integral method was used to predict the effect of pitch location and rate on the delay in quarter-chord separation. The wind-tunnel study involved the collection of time-varying pressure readings from 16 locations on an NACA 0015 airfoil that were subsequently used to determine lift, pressure-drag, and moment coefficients as functions of angle of attack for 140 test cases, covering 28 dynamic airspeed/pitch-rate/pitch-location combinations. Dynamic-stall effects of change (from steady flow) in the angle of attack at which separation occurs at the quarter chord (for comparison with the momentum-integral results), and change in the angle of attack at which stall occurs were extracted from these data and found to collapse best onto a non-dimensional pitch rate given by the chord times the pitch rate divided by two times the free-stream velocity. An adjusted non-dimensional rate formed by replacing one half the chord with the fraction of the chord corresponding to the pitch location was also examined and found not to be the proper non-dimensional variable for collapsing the data. The quarter-chord separation data compared favorably with the theoretical predictions.

Accelerated flow past a symmetric aerofoil: experiments and computations

2007 ◽  
Vol 591 ◽  
pp. 255-288 ◽  
Author(s):  
T. K. SENGUPTA ◽  
T. T. LIM ◽  
SHARANAPPA V. SAJJAN ◽  
S. GANESH ◽  
J. SORIA
Keyword(s):  
Reynolds Number ◽  
Flow Field ◽  
Angle Of Attack ◽  
Stream Velocity ◽  
Published Data ◽  
Navier Stokes Equation ◽  
Moment Coefficient ◽  
Flow Past ◽  
Naca 0015 ◽  
Accelerated Flow

Accelerated flow past a NACA 0015 aerofoil is investigated experimentally and computationally for Reynolds number Re = 7968 at an angle of attack α = 30°. Experiments are conducted in a specially designed piston-driven water tunnel capable of producing free-stream velocity with different ramp-type accelerations, and the DPIV technique is used to measure the resulting flow field past the aerofoil. Computations are also performed for other published data on flow past an NACA 0015 aerofoil in the range 5200 ≤ Re ≤ 35000, at different angles of attack. One of the motivations is to see if the salient features of the flow captured experimentally can be reproduced numerically. These computations to solve the incompressible Navier–Stokes equation are performed using high-accuracy compact schemes. Load and moment coefficient variations with time are obtained by solving the Poisson equation for the total pressure in the flow field. Results have also been analysed using the proper orthogonal decomposition technique to understand better the evolving vorticity field and its dependence on Reynolds number and angle of attack. An energy-based stability analysis is performed to understand unsteady flow separation.

scholarly journals Design and analysis of smoke flow visualization apparatus for wind tunnel

2021 ◽  
Vol 1206 (1) ◽  
pp. 012014
Author(s):  
D Raval ◽  
S V Jain ◽  
A M Acharii ◽  
K Ghosh
Keyword(s):  
Wind Tunnel ◽  
Flow Visualization ◽  
Operating Conditions ◽  
Design Flow ◽  
Column Height ◽  
Stream Velocity ◽  
Variable Frequency Drive ◽  
Power Inputs ◽  
Fan Speed ◽  
Naca 0015

Abstract In the present study, the design and analysis of smoke generator are done for the low-speed wind tunnel. The wind tunnel fan is fitted with the Variable Frequency Drive to produce the wind speed in the range of 3 to 32 m/s with fan speed of 150 to 1500 rpm. The design of smoke generator was done according to Preston Sweeting mist generator principle corresponding to the free stream velocity of 3 m/s. A controlled smoke generator consisting of kerosene reservoir, controlled heater, blower, liquid column height adjustment mechanism, valves etc. was designed and fabricated. The smoke generator produced the smoke at the rate of 154 cm3/s which was close to the design flow rate of 149 cm3/s. To supply the required quantity of smoke in the wind tunnel, the smoke rake of NACA 0010 profile was developed and installed in the rapid contraction section of the wind tunnel to achieve the streamlined flow. The parametric studies were done on the smoke generator at different power inputs and its effects were studied on smoke temperature, smoke discharge and boiling time of the kerosene. The flow visualization was carried out on NACA 0015 airfoil model and the images were captured to examine the flow physics around them under different operating conditions.

scholarly journals Effect of Flexible Flaps on Lift and Drag of Laminar Profile Flow

Energies ◽  
2020 ◽  
Vol 13 (5) ◽  
pp. 1077
Author(s):  
Artur Reiswich ◽  
Max Finster ◽  
Martin Heinrich ◽  
Rüdiger Schwarze
Keyword(s):  
Wind Tunnel ◽  
Critical Angle ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Force Balance ◽  
Airfoil Surface ◽  
Pressure Drag ◽  
Positive Effects ◽  
Trailing Edges ◽  
Lift And Drag

Experiments with elastic flaps applied on a common airfoil profile were performed to investigate positive effects on lift and drag coefficients. An NACA0020 profile was mounted on a force balance and placed in a wind tunnel. Elastic flaps were attached in rows at different positions on the upper profile surface. The Reynolds number of the flow based on the chord length of the profile is about 2 × 10 5 . The angle of attack is varied to identify the pre- and post-stall effects of the flaps. Polar diagrams are presented for different flap configurations to compare the effects of the flaps. The results showed that flaps generally increase the drag coefficient due to the additional skin friction and pressure drag. Furthermore, a significant increase of lift in the stall region was observed. The highest efficiency was obtained for the configuration with flaps at the leading and trailing edges of the profile. In this case, the critical angle was delayed and lift was increased in pre- and post-stall regions. This flap configuration was used in a gust simulation in the wind tunnel to model unsteady incoming flow at a critical angle of attack. This investigation showed that the flow separation at the critical angle was prevented. Additionally, smoke–wire experiments were performed for the stall region in order to visualize the flow around the airfoil. The averaged flow field results showed that the leading-edge flaps lean the flow more towards the airfoil surface and reduce the size of the separated region. This reduction improves the airfoil performance in the deep stall region.

Simulation of Flow Around a Static and Oscillating in Pitch NACA 0015 Airfoil Using URANS and DES

2004 ◽  
Author(s):  
J. Szydlowski ◽  
M. Costes
Keyword(s):  
Turbulent Flow ◽  
Numerical Simulations ◽  
Dynamic Stall ◽  
Cfd Applications ◽  
Naca 0015

This paper presents numerical simulations of the flow around a NACA 0015 airfoil at static and dynamic stall. The treatment of these configurations is a very challenging task for CFD applications. The turbulent flow around the static and in pitch oscillation airfoil is computed using different approaches: 2D RANS, 3D RANS and DES methodologies and with finer and finer meshes in order to try to reach a space converged solution. The main conclusion of the paper is that the prediction of static and all the more dynamic stall is not mature with present modeling capabilities.

Wind Tunnel Test on a Slowed Mach-Scaled Hingeless Rotor with Lift Compounding

2021 ◽  
Author(s):  
Shashank Maurya ◽  
Xing Wang ◽  
Inderjit Chopra
Keyword(s):  
Wind Tunnel ◽  
Reverse Flow ◽  
Wind Tunnel Test ◽  
Bending Moment ◽  
Flow Region ◽  
Dynamic Stall ◽  
Rolling Moment ◽  
Slowing Down ◽  
Slowed Rotor ◽  
Compressibility Effects

A single main rotor helicopter's maximum forward speed is limited due to the compressibility effects on the advancing side and reverse flow and dynamic stall on the retreating side. Compound helicopters can address these issues with a slowed rotor and lift compounding. There is a scarcity of test data on compound helicopters, and the present research focuses on a systematic wind tunnel test on lift compounding. Slowing down the rotor increases the advance ratio and, hence, the reverse flow region, which does not produce much lift. The lift is augmented with a wing on the retreating side. A hingeless rotor hub helps to balance the rolling moment with lift offset. Wind tunnel tests were carried out on this configuration up to advance ratios of 0.7 at two different wing incidence angles. Rotor performance, controls, blade structural loads, and hub vibratory loads were measured and compared with in-house comprehensive analysis, UMARC. A comparison between different wing incidences at constant total lift provided many insights into the lift compounding. It increased the vehicle efficiency and reduced peak-to-peak lag bending moment and in-plane 4/rev hub vibratory loads. The only trade-off was steady rotor hub loads and rolling moment at the wing root carried by the fuselage.

Optimization Calculation of Javelin Throwing Results

2014 ◽  
Vol 716-717 ◽  
pp. 764-766
Author(s):  
Min Jiang ◽  
Ji He Zhou
Keyword(s):  
Mathematical Model ◽  
Wind Tunnel ◽  
Initial Velocity ◽  
Angle Of Attack ◽  
Wind Tunnel Experiment ◽  
Aerodynamic Efficiency ◽  
Optimization Calculation ◽  
Computer Optimization ◽  
Necessary And Sufficient

On the basis of javelin wind tunnel experiment, we established mathematical model of javelin flight to conduct a computer optimization and got the conclusions. When the initial velocity is in the range of 25m/s-30m/s, the best throwing condition is: the throwing angle is 40°, the angle of attack is 11°. The javelin throwing condition is not zero angle of attack was necessary and sufficient for obtained aerodynamic efficiency.

Flight Characteristics of Flapping Wing Miniature Air Vehicles With “Figure-8” Spherical Motion

2009 ◽  
Author(s):  
Mohamed B. Trabia ◽  
Woosoon Yim ◽  
Zohaib Rehmat ◽  
Jesse Roll
Keyword(s):  
Mathematical Model ◽  
Wind Tunnel ◽  
Experimental Testing ◽  
Flapping Wing ◽  
Dynamic Stall ◽  
Wind Tunnel Testing ◽  
Servo Motor ◽  
Aerodynamic Forces ◽  
Flapping Motion ◽  
Spherical Motion

Hummingbirds and some insects exhibit “Figure-8” flapping motion that allows them to go through a variety of maneuvers including hovering. Understanding the flight characteristics of Figure-8 flapping motion can potentially yield the foundation of flapping wing UAVs that can experience similar maneuverability. In this paper, a mathematical model of the dynamic and aerodynamic forces associated with Figure-8 motion generated by a spherical four bar mechanism is developed. For validation, a FWMAV prototype with the wing attached to a coupler point and driven by a DC servo motor is created for experimental testing. Wind tunnel testing is conducted to determine the coefficients of flight and the effects of dynamic stall. The wing is driven at speeds up to 12.25 Hz with results compared to that of the model. The results indicate good correlation between mathematical model and experimental prototype.

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