Experimental Study of Flow Field on Upper Surface of NACA 0015 & NACA 2418 at DifferentAngle of Attack

Towsibur Rahman and Ariful Islam ShubhoDepartment of Aeronautical Engineering, Military Institute of Science and Technology, Dhaka, Bangladesh

Key words: Flow characteristics, flow-field, velocityvector, airfoil, angle of attack, flow velocity

Corresponding Author:Towsibur RahmanDepartment of Aeronautical Engineering, MilitaryInstitute of Science and Technology, Dhaka, Bangladesh

Page No.: 3741-3754Volume: 15, Issue 23, 2020ISSN: 1816-949xJournal of Engineering and Applied SciencesCopy Right: Medwell Publications

Abstract: The study of flow field over the upper surfaceof an airfoil has extensively been considered as a topic ofinterest in aerodynamic science. In this study, the flowfield over the upper surface of a symmetric airfoil and anasymmetric airfoil has been observed. For symmetricairfoil NACA 0015 has been used and for the asymmetricairfoil NACA 2418 has been employed. The method iscompletely usual but here we apply it to the conventionalproblem of the airfoils vertical to the wind. The study hasbeen carried out at Reynolds number, Re = 2*105. For thisReynolds number, the comparative analysis between theflow field of symmetric and asymmetric airfoils has beendone here. Main focus of this study is to determine thecharacteristics of flow field created over a symmetric andan asymmetric airfoil and find out the variation of thiseffect at different angles of attack. This studydifferentiates among the performances and sets upcharacteristics graphs which helps to understand the flowfields of two polar opposite objects in an easier way. Theslightest variation in angle of attack leads to a certainamount of impact which if studied can be helpful forvarious aerodynamic problems. So, an attempt has beenmade to find out this variation using closed loop windtunnel and two small scale airfoil models. Flow velocityhas been calculated by using total pressure and staticpressure with the help of Bernoulli’s principle. It issuggested to use advanced pressure sensors in order toachieve flow characteristics beneath the airfoils.

INTRODUCTION

Experimental study on airfoils has been conductedthroughout many years with the aim of differentiating theflow pattern over the upper surface of symmetric andasymmetric airfoil. It has been done for providing enoughinformation that will help solving aerodynamic problems

in future. Initially it was decided to do only experimentalstudy but with results of it further determination of flowdirection pattern was studied for better analysis of theresult.

Studies of airfoil flows have been motivated mostlyby efforts to avoid or reduce undesirable effects such asflutter, vibrations, buffeting, gust response and dynamic

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stall[1]. The future scope of the project is to simulate thedynamic stall phenomenon of any airfoil[2]. The study wasbased on the boundary layer concept of flow patternchange over airfoils.

There was the fact that the Wind Tunnel worked atlow velocity, resulting into variation of flow pattern fromthe original theoretical concept. This has motivated usinto making two new symmetric and asymmetric airfoilsand conduct the study on those rather using the existingones.

Flow over a symmetric airfoil, laminar flow patternchanges very slowly with the change of angle of attackwith a normal and regular change in velocity and pressureand there is a lesser chance of any wake region to becreated. However, if the angle of attack is increased thenthe wake region will commensurately increase. Whereas,flow over an asymmetric airfoil should be laminar in idealcondition but with the changes in angle of attack the flowturns into turbulent from laminar drastically. And in thiscase, velocity becomes negative due to turbulence.Negative velocity refers to change in velocity direction foradverse pressure. For this adverse pressure and change invelocity direction, there is always a chance of a wakeregion to be created in case of an asymmetric airfoil. Forsimilar rises of angle of attack wake region will begreater in the asymmetric airfoil condition than thesymmetric.

Experimental study of flow field on symmetric andasymmetric airfoil has been investigated in the followingarticles. Using linear structural theory, Nakano et al.[3]

investigated the flow field and tonal noise over the airfoil.Al-Garni et al.[4] studied the control of flow of airfoilusing experimental as well as numerical approach.Taslim et al.[5], through the simulation of an airfoilshowerhead hole design, performed experimental study ofimpingement on roughened airfoil leading-edge walls with film holes. Tian et al.[6] applied an experimentalstudy to identify the performance of pitching airfoil.Chong and Joseph[7] studied and extended airfoilinstability. Graziani et al.[8] performed an experimentalstudy of end-wall and surface of the airfoil.Kotsonis et al.[9] investigated circulation control of airfoilusing plasma actuators. Genc[10] conducted Anexperimental approach on aerodynamics of NACA2415airfoil at low Re numbers. O’meara and Mueller[11]

employed an airfoil at Reynolds number to identify thelaminar separate bubble characteristics. Lai and Platzer[12]

studied the characteristics of jet by applying plungingairfoil.

Research in the field of velocity and flow is alsoabundant. Raffel et al.[13] studied unsteady flow velocityfield over an airfoil. Suzuki et al.[14] investigated theunsteady PTV velocity field past an airfoil. Ladopoulos[15]

obtained experimental results for inviscid flow-fields ofunsteady airfoils. Amitay and Glezer[16] carried out

experiments Controlled transients of flow re-attachmentover stalled airfoils. Rojratsirikul et al.[17] provided thevelocity field for membrane airfoil. Towsibur and Islam[18]

identified flow characteristics and velocity field ofcylinder which can be applied in airfoil. Garcia-Sagradoand Hynes[19] analyzed velocity of flow near the trailingedge of naca 0012 airfoil. Siauw et al.[20] investigatedvelocity field near wake region of naca 0015 airfoil.Chen and Ho[21] measured the two-dimensional flow fieldvelocity. Lee and Jang[22] studied the mean velocity fieldsof the near wake behind the smooth and MRF-coveredairfoils. Ol et al.[23] identified laminar separation bubbleseparation by the velocities in the free-stream which by virtue of the stagnant fluid. Derksen et al.[24] employedimage velocimetry (PIV) technique for the velocitymeasurements.

The main targets behind the study were to modifyYaw meter and Pitot tube, build one symmetric and oneasymmetric airfoil, understand the flow pattern overdifferent airfoils, gather experimental data for new airfoilsand finally, compare the experimental data withtheoretical concept.

This study is for finding out the noticeabledifferences between the flow properties of symmetric andasymmetric airfoils which when compared to thetheoretical concept will provide a better knowledge aboutthe aerodynamic properties of various airfoils to solvefuture problems related to this topic. The objectivesbehind the study are to:

C Modification of Yaw meter and Pitot tubeC Build one symmetric and one asymmetric airfoilC Understand the flow pattern over different airfoilsC Gather experimental data for new airfoilsC Compare the experimental result with the theoretical

concept

MATERIALS AND METHODS

FormulationBernoulli’s equation: It is the relation between pressureand velocity in an inviscid, incompressible flow. Theequation is:

1P+ ρV2 = const

2

This equation is called Bernoulli’s equation.Bernoulli’s equation is probably the most famousequation in fluid dynamics. Considering the x componentof momentum equation for inviscid flow with no bodyforces, we get:

(1)

Du pρ = -

Dt xu u u u p

ρ +ρu +ρv +ρw = -t x y z x

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For steady flow, Equation 1 is then written as:ˆ/ t 0.

u u u 1 pu +v +w = -

x y z ρ x

Multiplying equation by dx:

(2)u u u 1 p

u dx+v dx+w dx = - dxx y z ρ x

Consider the flow along a streamline inthree-dimensional space. In particular, substituting:

udz-wdx = 0

vdx-udy = 0

Into Eq. 2, we get:

(3)u u u 1 p

u dx+v dx+w dx = - dxx y z ρ x

(4)u u u 1 p

u dx+ dx+ dx = - dxx y z ρ x

Recall from calculus that given a function u = u(x, y,z), the differential of u is:

u u udu = dx+ dx+ dx

x y z

This is exactly the term in parentheses in Eq. 4.Hence, Eq. 4 is written as:

(5)

2

1 pudu = - dx

ρ x

or:

1 1 pd u = - dx

2 ρ x

In a similar fashion, starting from the y component ofthe momentum equation specializing to an inviscid,steady flow and applying the result to flow along astreamline, we have:

(6) 21 1 pd v = - dy

2 ρ y

Similarly, from the z component of the momentumequation, we obtain:

(7) 21 1 pd w = - dz

2 ρ z

Adding Eq. 6 through 7 yields:

(8) 2 2 21 1 p 1 p 1 pd u +v +w = - dx+ dy+ dz

2 ρ x ρ y ρ z

(9) 2 2 2 2u +v +w = V

(10)u u u

dx+ dx+ dx = dpx y z

(11)

21 pV = -

2 ρ

u u udx+ dx+ dx

x y z

dp = -ρVdV

Equation 11 is called Euler’s equation. It applies toan inviscid flow with no body forces and it relates thechange in velocity along a streamline dV to the change inpressure dp along the same streamline. Equation 12 takeson a very special and important form for incompressibleflow. In such a case, ρ = constant and Eq. 12 can be easilyintegrated between any two points 1 and 2 along astreamline. From Eq. 12 with ρ = constant, we have:

(12)

2 2

1 1

P V

P V

2 22 1

2 1

2 21 1 2 2

dp = dV

V VP -P = -ρ -

2 2

1 1P + ρV = P + ρV

2 2

Equation 12 is Bernoulli’s equation which relates P1

and V1 at point 1 on a streamline to P2 and V2 at anotherpoint 2 on the same streamline. Equation 13 can also bewritten as:

(13)2 constant; along a stre1

P+ ρ amV2

line

Relation between velocity and manometric height:From Bernoulli’s equation, we get:

2 21 1 2 2

1 2

P V P V+ +z = + +z

γ 2g γ 2g

In case of pitot tube or yaw meter, Z1 = Z2 and V2 =0. So, we get:

(14)

22 1 1

22 1 1

22 1 1

2 2 11

P P V= +

γ γ 2g

P P V- =

γ γ 2g

P -P V=

γ 2g

P -PV = 2g×

γ

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J. Eng. Applied Sci., 15 (23): 3741-3754, 2020

We know, P = γh. From Eq. 14, we get:

(15)21V = 2gΔh

(16)

a a k k

ka k

a

a k

γ h = γ h

γh = h

γ

h = 0.64h

Substituting Eq. 16 into Eq. 15:

(17)

21 a

21 k

1 k

1 k

V = 2g×h

V = 12.54×h

V = 12.54× L sin30

V = 2.5× L

This is the desired relation between velocity andmanometric height and used to calculate static velocityduring this study[25].

Three tube probe equation: Provided that flow directionremains within a limited range of pitch, three-tube probeconsisting of a forward facing pitot tube and two inclinedside tubes can be used to determine wind speed inaddition to flow direction in one plane:

C P1,3 = Pressure measured by inclined tubes of yawmeter

C P2 = Pressure measured by the central tube of yawmeter

A number of three tube probes which could be usedwith a method in which the probe is rotated in yaw aboutits apex until the side tube register the same pressure(Fig. 1). Total pressure is then read directly from thecenter tube, whilst kinetic pressure is taken as beingproportional to the difference between the center tubepressure and that of the side tubes (P1,2). To determineflow angle calibration process is required and thecalibration process uses the following simple straight lineequation:

(18)tanθ = mΨ+c

1 3

1 32

P -Pψ =

P +PP -

2

Using the calibration data, tan θ versus Ψ can beplotted. From the graph the slope (m) and constant (c) canbe determined. As the value of can be obtained fromexperiment, flow direction θ can be easily found from Eq. 18.

Fig. 1: Three tube probe

Fig. 2: Top wall

Experimental setup: The experiment required for aspecific orientation and design for measurement system.So, test section was designed and modified viewing thispurpose. This part encloses the whole experimental setupfor this specific experimental study.

Test section: The test section was the most vital part ofthe modification of the experimental setup. The testmodels were set in the test section with the help ofmounting system. Eventually, the top wall of the testsection was modified in a manner, so that, pressure,velocity and flow direction can be found over each pointof the test model and also upstream and downstream ofthe test model. The material used was plexi-glass sheet of5 mm thickness.

Top wall: The top wall was modified for accommodatingtraverse mechanism at upstream position at 60 mm fromfront of test section and a hole was created for static portwhich is about 36 mm from the front of the test section. Agroup of 7 mm breadth and 170 mm length was created at196 mm from the front of the test section. For the creationof this group the yaw meter can be moved all over the testmodel and also at upstream and downstream of the airfoil(Fig. 2).

Slider: On the top wall a slider was used at which threeholes were created at 90, 70 and 30 mm, respectivelyfrom the front of the slider for fitting yaw meter. The yaw

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2

3

1

J. Eng. Applied Sci., 15 (23): 3741-3754, 2020

0 0.5 1.0 1.5 2.0 2.5 3.0

0.7

0.6

0.5

0.4

0.3

0.2

0.1

0

f'a

Ψ

Fig. 3: Slider

Fig. 4: Traverse mechanism

meter is fitted on the slider with the help of traversemechanism. The slider allows the yaw meter to move overthe airfoil surface in x- direction and also at upstream anddownstream of the airfoil. The diameter of the holes foraccommodating the yaw meter is 7 mm (Fig. 3).

Yaw meter: Yaw meter was made using threeconventional needles of syringe with 1.2 mm diameter.The needles were attached with a stainless steel rodof 6 mm diameter. Each needle is connected with a tubeof 1.5 mm diameter which is passed through the rod andcan be connected to the manometer.

Traverse mechanism: Traverse mechanism is used tomove the yaw meter or pitot tube at different positions inone specific direction. Two traverse mechanisms wereused in the experiment. One for yaw meter and anotherfor pitot tube. The traverse mechanism has different parts(Fig. 4).

Rotating screw: The main part which allows the yawmeter or pitot tube for moving is a screw. Whileaccording to vector multiplication, rotating screws are totravel in direction perpendicular to rotation, this screw

Fig. 5: Calibration of yaw meter

stays still but enables adjacent sliding plane to travel forit. The screw head while rotated in clockwise directionmoves up and for counter-clockwise direction it movesdown.

Scale: A millimeter scale is used to measure travelingdistance.

Indicator: Adjacent to sliding plane, this indicator showsthe position on scale.

Sliding plane: It houses the yaw meter or pitot tube atholder. And it is made so that it reacts to screw rotationand moves up and down. The indicator is also reliedon it.

Yaw meter calibration: Yaw meter was calibrated in aswirling flow field within engineering tolerance todetermine the value of m and c from the equation statedbelow[26, 27]:

tanθ = mΨ+c

The values from the table were used to plot the graph(Table 1 and Fig. 5). From the graph the m and c weredetermined for further use in calculation of velocitydirections. From graph:

m = 0.295

C = -0.1

Neutral condition of setup: The neutral condition of thesetup was tested for obtaining initial condition andcorrection of measured values in future. Thus, it is abackground of which the study deals with. Neutralenvironment is defined by test section without any modelor obstruction or any other means which might influenceairflow inside it.

Experimental approachConnection of pitot tube: A pipe from pitot tube and apipe from static port were connected to a manometerwhich is inclined at a particular angle. It was ensured thatfor that particular angle, all the initial reading of

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Table 1: Yaw meter calibration tableα tanα P1 P2 P3 Clockwise Anticlockwise 0 0 14.2 9 27 0.34 0.345 0.088 11.75 10.5 26.5 0.08 0.610 0.1763 9.25 12.25 26.75 -0.1875 0.8715 0.268 6.5 14.5 26.5 -0.5 1.182520 0.364 3 16.75 25.25 -0.89 1.5725 0.466 0.75 18.5 22 -1.43 2.1130 0.577 -0.25 19.5 18.5 -2.11 2.79

connected tube for ambient pressure are same. Thus, therewill be no variation for initial condition. Upstream pitottube was kept at middle position of test section so that theconsistent value of upstream total pressure can beobtained from the inclined manometer. The angle ofinclined manometer was at 30°, so that, initial positionalways remains same. Before stepping for furtherprocedures, it was checked that the pitot tube as well aspipes are free from any obstructions such as dust orblockage due to glue during connection between twopipes of different diameter.

Connection of yaw meter: Yaw meter consists of threetubes. Each tube of yaw meter was connected to aseparate inclined U tube manometer and the other tube ofthe manometer was kept open to atmospheric pressure.This inclined manometer was also kept at 30° duringexperiment. Before stepping for further procedures, it waschecked that the yaw meter as well as pipes are free fromany obstructions such as dust or blockage due to glueduring connection between two pipes of differentdiameter or connection with the nipple. Before experimentit was also checked that the pipe fittings are free from anytype of leakage (Fig. 6).

Reynolds number set-up: The closed loop wind tunnelat which the experiment was conducted has two motors ofsame power and same rpm. Each of them can separatelysupply flow about 10-12 m secG1. When only one motorwas turned on the Reynolds number becomes 1.45*105

and when two motors are turned on the Reynolds numberbecomes 2*105. By turning on both the motors we havedone our experiment.

NACA 2418 airfoil set-up: NACA 2418 was firstmounted at 0° AOA for the experiment and total andstatic pressure were measured. Then traverse mechanismwere fitted at different position of the slider separatelyand pressures at three tubes of yaw meter were measuredat various points over the airfoil and at upstream of theairfoil. The same process was repeated for differenty-position that means at different heights over the airfoil.And the whole above process was done for the Reynoldsnumber 2*105.

The AOA of the airfoil was gradually changed from0° to different angles varied as 5, 10 and 15° and the

Fig. 6: Connection of pitot tube and yaw meter

Fig. 7: NACA 2418 Airfoil Set-up at 0° AOA

experiment was continued similarly as done for 0° AOA.Demonstration of NACA 2418 airfoil set-up at 0° AOA isgiven in Fig. 7

NACA 0015 airfoil set-up: NACA 0015 was firstmounted at 0° AOA for the experiment and total andstatic pressure were measured.

Then traverse mechanism were fitted at differentposition of the slider separately and pressures at threetubes of yaw meter were measured at various points overthe airfoil and also at upstream of the airfoil. The sameprocess was repeated for three different y-position thatmeans at three different heights over the airfoil. And thewhole above process was repeated and done at Reynoldsnumber 2*105.

The AOA of the airfoil was gradually changed from0° to different angles varied as 5°, 10°, 15° and theexperiment was continued similarly as done for 0° AOA.Demonstration of NACA 0015 airfoil set-up at 0° AOA isgiven in Fig. 8.

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Fig. 8: NACA 0015 airfoil set-up at 0° AOA

Fig. 9: Inclined manometer

Measuring system: The overall system consists andrequires some measuring system as stated below:

Inclined manometer: For low levels of deflections,manometers are used to measure pressure of fluids.Traditional variants include single tube, multi tube,U-tube, vertical, inverted, inclined ones, etc. The mostpopular use for manometer is probably measuring staticand total pressure fluid (Fig. 9).

To conduct the experiment four U-tube manometersare required as yaw meter has three probes and for thestatic pressure value. So, four U-tube manometers wereset up on a wooden box using glass tube, pipe fittings,measuring scales, etc. Wooden blocks were used to keepthe manometers inclined at different angles. As the studyused a U-tube manometer with inclination the differenceof ambient to manometer pressure is:

mP = ρ gh

While inclined, manometer pressure is given by:

mP = ρ gLsinθ

Where:L = The lengthθ = The angle of inclination

Another inclined manometer was made to measurethe total pressure and static pressure at upstream. From

Fig. 10: Inclined manometer (Side view)

the difference of total and static pressure, we can calculatethe flow velocity of the wind tunnel and also the Reynoldsnumber (Fig. 10).

Pitot tube: Most common device to measure air flowpressure. A pitot tube measures total pressure consistingof both its static and dynamic variant. With the help of astatic pressure measuring unit, dynamic pressure thusflow velocity can be obtained:

2 2A A B B

21 1

o 11

o

1 1P + ρV = P + V

2 21

P + ρV2

2 PV =

P

-P

ρ

In short, total pressure is measured from pitot tubeand static pressure from static port, the difference betweenthem gives flow velocity measurement eventually. Here,ρV2 gives dynamic pressure variant of flow. This theoryholds for only incompressible flow and as our specificwind tunnel supplies us with low speed airflow this ismost convenient means to use.

RESULTS

Symmetric airfoil: In horizontal axis velocity profile ispresented as well as in vertical axis yaw meter height isshown. Different angle of attack is used to comprehendthe flow velocity profile over the airfoil.

Overall, velocity profile range from approximately0.097 m secG1 to about 1.13 m secG1. Yaw meter height isfrom 0-0.35 cm. As the flow is not influenced in the upperportion of the airfoil, flow velocity remains constant in theupper portion of the test section. However, in the origin ormiddle portion flow velocity greatly deflected by thepresence of the substance. Hence, the flow velocityreduces and create turbulence. In the airfoil, as flowcontacts across the leading edge to trailing edge flowturns laminar into turbulent. By the boundary layer theory,flow decreases in the wall and that’s why flow is lowestin the surface or boundary of the airfoil.

3747

P2

0

P1 Inclined tube

L

H = L sin θ θ

J. Eng. Applied Sci., 15 (23): 3741-3754, 2020

Fig. 11(a-d): Velocity profile over the upper surface of NACA 0015 airfoil at Re = 2*105, (a) AOA0°, (b) AOA 5°, (c)AOA10° and (d) AOA 15°

In Fig. 11a, flow is constant in the upper portion ofthe airfoil is steady while in lower portion it significantlychanges. When x/c is 0, flow strikes the origin of theairfoil while in upper portion flow velocity remainsconstant. It is generally known as stagnation point whereflow pressure is maximum but flow velocity is minimum(approximately 0 m secG1). Then with the increment of x/c(horizontal distance) flow velocity decreases. When x/c isequivalent to 1, flow velocity is >0.7 m secG1 but<0.8 m secG1 which is the lowest velocity of the 0 angleof attack position of the airfoil. At first, flow in the uppersurface of the airfoil is laminar but then it alters intoturbulent as par as the boundary layer theory.

The stagnation point progressively movesdownstream of the leading edge over the bottom surfaceof the airfoil as AOA (5° angle of attack) is increased.For that reason in Fig. 11b, at x/c = 0, flow velocity

slightly greater than the 0° AOA. In upper portion likebefore velocity is in the steady condition. At x/c = 1, flowvelocity is >0.6 m secG1 but <0.7 m secG1.

Angle of attack further increases in Fig. 11c andthat’s why stagnation point again moves ahead. Thisincrement of angle of attack makes flow laminar toturbulent faster. Initial velocity is around 0.97 m secG1.Afterwards velocities keep reducing as well as it isbecoming turbulent for the existence wake region justbehind the separation of the flow. Flow becomes chaoticfrom leading edge to trailing edge. At the 1 cm of theposition of x/c, flow velocity is about 0.5 m secG1 whichis lesser than previously 5° angle of attack.

In Fig. 11d, flow velocity greatly reduces as hereangle of attack further increases which is 15°. Flow fieldis attached over the top surface of the airfoil. However, asAOA is again increased, massive flow-field separation

3748

(a0.25

0.20

0.15

0.10

0.05

0

y/c

0.4

(c0.35

0.30

0.25

0.20

0.15

0.10

0.05

0

y/c

0.4

a)

0.5 0.6

c)

0.5 0.6 0

xc/ = 0 xc/ = 0.2 xc/ = 0.4 xc/ = 0.6 xc/ = 0.8

xc/ = 1.0

0.7 0.8 0.9

u/U

0.7 0.8 0.9

u/U

1.0 1.1

y/c

1.0 1.1

y/c

(b) 0.30

0.25

0.20

0.15

0.10

0.05

0 0.4 0.5

(d)

0.25

0.20

0.15

0.10

0.05

0

y/c

0 0.2

0.6 0.7 0.8

u/U

0.4 0.6

u/U

8 0.9 1.0

0.8 1.0

1.1

1.2

J. Eng. Applied Sci., 15 (23): 3741-3754, 2020

Fig. 12(a-d): Velocity profile over the upper surface of NACA 2418 airfoil at Re = 2*105, (a) AOA 0°, (b) AOA 5°, (c)AOA 10° and (d) AOA 15°

occurs against the top surface. The type of stallphenomenon occurs is leading edge stall; it ischaracteristics of relatively thin airfoils with thicknessratios between 10 and 16% of the chord length. As seenabove, flow separation takes place rather suddenly and abruptly over the entire top surface of the airfoil withthe origin of this separation occurring at the leadingedge.

Asymmetric airfoil: NACA 2418 is fabricated andemployed to determine the velocity alongside flowcharacteristics over the upper surface of the airfoil. Hereflow-field get reduced with the horizontal distance of theairfoil. At the end for the generation of the huge wakevelocity greatly changed. It turns into turbulent for theadverse gradient pressure.

Overall, flow velocity is steady across the upperportion of the test section because there is no influence ofany object. But when the flow senses the existence of theairfoil, it subsides its speed. That’s why flow here in theupper portion keeps reducing.

In Fig. 12a where x/c is equal to zero, flow isconstant in the upper region. Yet, flow does not have anymark for the airfoil. It has not any influence for x/c = 0.2,0.4 as well. At x/c = 0.6, in the surface of the cylinder,flow velocity is around 1.009 m secG1. As yaw metermove ahead, speed again decreased (x/c = 0.8). When x/cis 1, flow speed subsides most. There, velocity is>0.75 m secG1 but <8 m secG1. At 0° AOA flowgeneration from laminar to turbulent is quite slower thanother AOA position of the airfoil.

Angle of attack is 5° in Fig. 12b where flowtransforms laminar to turbulent quickly than 0° angle of

3749

2

0.2

0.1

0.1

0.0

y/c

0.2

0.2

0.

0.

0.0

y/c

25

20

15

10

05

0

0.60 0.65 0.70 0

(a)

xc/ = 0 xc/ = 0.2 xc/ = 0.4 xc/ = 0.6 xc/ = 0.8

xc/ = 1.0

(c) 25

20

15

10

05

0

0 0.2 0

0.75 0.80 0.85 0.

u/U

0.4 0.6 0

u/U

.90 0.95 1.00 1.0

0.8 1.0 1.

5

(b) 0.25

0.20

0.15

0.10

0.05

0

y/c

0.4

2

y/c

-0.2

(d) 0.

0.

0.

0.

0.

0.5 0.6 0.7

u

0 0.2 0.4

25

20

15

10

05

0

0.8 0.9

u/U

0.6 0.8 u/U

1.0 1.1

1.0 1.2

J. Eng. Applied Sci., 15 (23): 3741-3754, 2020

Fig. 13(a-d): Velocity vector field of NACA 0015, (a) AOA 0°, (b) AOA 5°, (c) AOA10° and (d) AOA15°

attack. With the reduction of the distance by 0.2 flowvelocity greatly reduces than previously. End of thesection of airfoil (x/c = 1), speed of the flow becomesapproximately 0.65 m secG1. Figure 12c and 12d flowagain subsides. Lowest speed for AOA 10° is about0.3 m secG1 and for AOA 15° is around 0 m secG1. AOA15°, flow heavily affected by the turbulence. For theadverse pressure gradient flow converts laminar toturbulent quickly than any other AOA position.

First of all, pressure is highest for the certain positionof the airfoil. Therefore, velocity is lowest in that region.Afterwards, pressure reduces with the increment of thehorizontal position with respect to the airfoil. Hence,velocity also decreases with the parallel incrementposition of the airfoil.

Verification: A MATLAB code is employed to generategraphs at different AOA for the two airfoils for furtherreliability on our observation results. Quiver plot is usedto analyze the velocity vector field.

By using manometer reading from yaw meter (P1, P2,P3), Yaw meter position (y/c) in y direction, static/wall Pressure (Ps) and flow angle flow velocity resultantdirection is determined which is:

-1θ = tan (mΨ+c)

Then total magnitude velocity is evaluated by u/U.Afterwards V*cosθ and V*sinθ is identified to understandthe overall flow-field over the airfoil.

With the increment of the angle of attack flowbecomes chaotic and flow velocity magnitude decreasessignificantly. At AOA 15 in Fig. 13d, flow velocityhugely affected and here turbulence region is greater thanany other AOA airfoil turbulence region. In the upperportion of the test section, flow is in the same directionand steady for AOA 0° (Fig. 13a). Then flow velocitychanges where x/c = 1. Here flow reduces in the surfaceof the airfoil. At the position of 10° angle of attack flow

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0

0

0

0

0

-0

y/c

0

0

0

0

0

-0

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(a) 0.25

0.20

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0

0.05

0 0.2

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2

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y/c

0 0.2 0.4 0u0.6 0.8 1.u/U

0 1.2

J. Eng. Applied Sci., 15 (23): 3741-3754, 2020

Fig. 14(a-c): Velocity vector field of NACA 2418, (a) AOA0°, (b) AOA 5° and (c) AOA 10°

changes not only in x/c = 1 but also in x/c = 0.8. Flowdirection also changes a lot for the presence of the wakein the surface of the airfoil.

The above MATLAB generated graphs show that athigher AOA flow changes from laminar to turbulent. Aswe did not take readings after the trailing edge so wakeregion did not appear in the graphs. But flow changinginto turbulent phenomenon was observed. As few of thevalues came in negative with which we could notcalculate the velocity, the graph at 10° AOA (Fig. 14c) forthe asymmetric airfoil could not be achieved. This showedthat flow separation occurs much earlier on theasymmetric airfoil than on the symmetric airfoil and theasymmetric airfoil was already at stall at 15° AOA.

DISCUSSION

As per our experimentation and results variouscharacteristics of both the symmetric and asymmetricairfoils is observed. It is discussed in a broad way in the

following context. At 0° AOA, laminar airflowcharacteristic has seen over the upstream of the symmetricairfoil. This corresponds to the actual phenomenon thathas already been established and discussed earlier in thispaper. As, with the increment of AOA and airflowoccurred at the upstream of the airfoil, there is a transitionfrom laminar to turbulent flow. At 10° and 15° mostchange occurring is seen as well as could observeturbulent region characteristics. Values are taken in thex-direction at only 6 points which has assisted ourobservations results. This assists to pinpoint the exactlocation in the x-direction at which the transition occurs.There has been an observation of the similar phenomenonof laminar flow changing into turbulent flow as per theincrease in AOA and as per the change in x/c positionnear the trailing edge. Taking values after the trailingedge area would have provided extended observation onwake characteristics.

By the observation of flow characteristics of thesymmetric airfoil, transition of airflow from laminar to

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(a) (b) 0.25

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0 0.2 0.4 0.6 0.8 1.0 1.2 u/U

0 0.2 0.4 0.6 0.8 1.0 1.2 u/U

(c)

0 0.2 0.4 0.6 0.8 1.0 1.2 u/U

0.25

0.20

0.15

0.10

0.05

0

-0.05

y/c

J. Eng. Applied Sci., 15 (23): 3741-3754, 2020

Fig. 15(a-f): Asymmetric airfoil and symmetric airfoil, (a) Velocity profile on the upper surface of naca 2418 at 10° andat x/c = 1, (b) Velocity profile on the upper surface of naca 0015 at 10° at x/c = 1, (c) Velocity profile onthe upper surface of naca 2418 at 15° and at x/c = 0.8, (d) Velocity profile on the upper surface of naca0015 at 15° and at x/c = 0.8, (e) Velocity profile on the upper surface of naca 2418 at 15° and at x/c = 1and (f) Velocity profile on the upper surface of naca 0015 at 15° and at x/c = 1

turbulent is encountered with the increase in AOA and x/cposition movement towards the trailing edge. Now if wetry to compare both the airfoils we need to consider theirchange in airflow with change in AOA and change of x/cposition. For the asymmetric airfoil at 0° AOA it wasclearly seen that laminar flow is occurring almosteverywhere from leading edge to trailing edge of theairfoil. But there was sudden change as we changed theAOA from 5-10° which was not too much significant forthe symmetric airfoil. And at 15°, we could observe theutmost turbulent region and almost wake characteristicswere achieved as few values near the trailing edge came

in negative which could not be used in formulas tocalculate the velocity ratio. However, our observationresult for the greater AOA was more accurate than theresult of 0° AOA. Because at 0° and at x/c position 0, y/cvalue could not be taken at the boundary layer due to theyaw meter probe’s lack of precision. At position x/c = 0and at AOA 0°, y/c position of the yaw meter probe isabout 3 cm above the boundary layer which made usincapable of achieving the true characteristics at thatposition. As a result of it some unusual characteristicsoccurred for the x/c = 0 position in both the symmetricand asymmetric airfoils (Fig. 15).

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ues

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ues

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J. Eng. Applied Sci., 15 (23): 3741-3754, 2020

The comparison table shows how the flow propertiesvary for each of the airfoils with increase in AOA andchange of x/c position (Fig. 15).

CONCLUSION

The difference in flow characteristics over asymmetric and asymmetric airfoil is carried out toexperience and compare with the actual phenomena.NACA 0015 is used as the symmetric airfoil and NACA2418 has applied as the asymmetric airfoil. Small scaleairfoil model is employed and two dimensional flow perunit width is developed over the model to allow themeasurement of airfoil properties instead of a finite wing.Bernoulli’s equation is applied to illustrate relationbetween velocity, manometric height and three-tube probeequation to determine the flow velocity and direction overthe airfoils.

In the experiment, flow property only at Reynoldsnumber 2*105 is evaluated. At lower Reynolds number,more accurate flow characteristics can be found out.Velocities over the airfoil has measured at 1.5 cm intervalin the x-direction and at various positions in y-directionfor each of the points to plot the velocity profile graphs.Results came similar to the actual phenomena but at theleading edge flow characteristic at 0° AOA is notidentified due to lack of precision of the three tube proberod. By modifying the test section and using advancedpressure sensors flow characteristics beneath the airfoilcan be achieved. Overall our results were close enough tothe actual flow characteristics of both the airfoils as wellas may be this can be used for further study in this field.

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