A Visualization Low Speed Subsonic Wind Tunnel Design and ...

Journal of Mechanical Engineering Research and Developments

ISSN: 1024-1752

CODEN: JERDFO

Vol. 43, No. 5, pp. 17-28

Published Year 2020

17

A Visualization Low Speed Subsonic Wind Tunnel Design and

Construction for Laboratory Application and Uses

Ahmed Adnan Shandookh, Salman H. Omran & Laith Jaafer Habeeb*

Mechanical Engineering Department, Energy and Renewable Energies Technology Center & Training and

Workshop Center, University of Technology, Baghdad, Iraq

*Corresponding Author Email: [email protected]

ABSTRACT: The aim of the research is to design and construct a low speed small model of a sub-sonic wind

tunnel under standard dimensions for laboratory application and purposes which is able to handle models of

different types and/or sections, as well as different wing sections, and calculating the lifting, drag and moment

coefficients around these sections through calculating the pressure coefficient. The model was constructed using

high stainless-steel sheets that can withstand high stress levels. The inspection area is also equipped with glass

panels so that we can observe the air flow around the different sections. The section is also manufactured with a

hole above and below the section. The low speed small wind tunnel is also equipped with a digital manometer for

calculating the rate of pressure variation around the sections as well as digital flowmeter that calculating the speed

of flow. The wind tunnel is also equipped with a high-efficiency fan at one high speed and has been modified and

re-controlled so that five different speeds can be applied to the device. The preliminary results were promising and

by comparing those results with theoretical calculations by applying more than one specialized program, the results

were found to be significant and the laboratory could be adopted.

KEYWORDS: Wind tunnel; Pressure distribution; Lift

INTRODUCTION

Simply, Wind tunnels are defined as an instrument for airflow stream calculations and tests. It consists basically

from three parts. The first (Intake) is usually diverging converging tunnel, the second one is straight tunnel in

almost all types and finally the last part (Exit) is converging diverging nozzle tunnel. Wind tunnels also used to

learn more about aircraft, ships and vehicles models in order to understand the basic fundamentals of aerodynamics

and its characteristics coefficients. Where it helps to improves these models as well as reduce fuel conception by

enhancing the aerodynamic shape of it and hence reduce the friction over its surface. In some application this will

also increase the estimation of fatigue life of the material used. Some of these wind tunnels are so big that it could

hold a large model as well as vehicles. In these wind tunnel air or fluid were moves over and around these models,

then, some source of detecting sensors was used to estimate and measure the required values.

Basically, a huge powerful motor with high performance fan or impellers uses to generate the air/fluid inside the

wind tunnel. While, the model that need to be tested is fixed in the second part tunnel (medial one usually), with

a specific position and/or angles in some models, then with the flow boundary moves over these models the

calculation begins. Although, the models could be big or small one but also it could be some little part of a model

or vehicle. Even in some types a very small model. Many of these wind tunnels was manually handling, but some

of these wind tunnels was so improved and complicated that uses a high technology for calculation. The models

even if it is fixed in most of the cases but some of these wind tunnels where able to handle with moving models

these are the most complicated CFD tests. Some studies concentrated on the structure of the models, while others

were dell with the flow around these models. In some of these wind tunnels a smoke and/or dye can be mixed with

the air to enhance the flow generation. During the test the speed of the air/fluid is changes from subsonic to

transonic even in some wind tunnel a supersonic speed. In the present day, there are a lot of different types, speeds,

application and volumes of wind tunnels.

WIND TUNNEL APPLICATION

The application of wind tunnels is numerous, but basically its used for design improvement and/or enhancing,

most of these applications were for civilian uses, but, in some cases, it could be used for military application also,

especially in military ships and aircrafts, through the application of these wind tunnels many improvement,

mailto:[email protected]

A Visualization Low Speed Subsonic Wind Tunnel Design and Construction for Laboratory Application and Uses

18

enhancement, re-design, design accessories and so were applied for many models of ships, aircrafts and vehicles.

Since, 19th century improvements and enhancements were applied till our present day. It began with a simple low

speed, small test section and man handle devise, while the modern one was too complicated, high speed and

completely computerized wind tunnels. Basically, wind tunnels help to develop, improve and enhance aircrafts

types and models through the numerus studies and researches that concentrates upon studying the fundamentals

of aerodynamic characteristics and coefficients which leads finally to improve flying theories.

The use of the wind tunnel has evolved in most engineering fields. We mention some of these areas (See Figure

(1)):

1. It is used in the field of civil engineering in terms of testing buildings and affected by wind speed.

2. It is used in military fields, where combat aircraft, missiles, and the like are tested.

3. Used in the field of civil aircraft, where all parts of the plane are inspected such as wing, tail.

4. Used in the measurement of aerodynamic forces generated on cars of all types, small and large and even

racing cars.

Figure 1. Some Applications of Wind tunnel.

BASIC THEORY OF WIND TUNNEL

There are many ways to calculate and investigate the basic aerodynamic characteristic coefficients principals like

CL, CD and the Cm only from the estimated flow pressure distribution that takes over the airfoil wing model section

of any model, in this research NACA 0015 will be used for these calculation, this could be made experimentally

in a specific wind-tunnel, (Figure-1), represents some kind of section of a selected aerofoil wing model at some

incidence angle to the air fluid airstream, which could be in some cases assume to arrive from the left to right

positions, at a suitable selected airflow speed velocity of (V). A drawn of the selected axes (ox) and (oz) were

made through the nose of the airfoil model section and they are respectively represented parallelly and

perpendicularly to the chord line of the wing section model. The main chord of the airfoil model wing section was

denoted by (c). While (z1) and (z2) were the coordinates of the highest and lowest points of the surface section

position respectively, (see Figure-2).

The aerofoil wing section model with its surface is to be consider as a small thickness plate material, even if it

might be or might not be solid, also it could be assumed as perfectly rigid body with the initial inside pressure

(Po)[Pa], that uniformly at this point, since, the estimated static pressure of the un-disturbed airstream of the wing

section model is premise, where, the arrived pressure were assumed to be integrated all around upper and lower of

the surface of the wing section model, and also, the integral of the calculation of a uniformly pressure occurs over

the fixed surface is assumed to be equal to zero, even if whatever its magnitude was. If we take a little element of

this aerofoil of a unity length for span-wise, and if considering the real affecting forces on a relatively very small

piece of element, with a determinant length of (s )[mm], relative to the aerofoil wing section model surface. Then,

a perpendicular force of these element selected, is said to be completely a composed of (s Po ) out wards, and (

sP ) in wards, and, as well as some forces ofs Po)-(P .


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Finally, this force might be resolved into two components ( x ) and ( z ) parallelly and respectively acting to

the (ox) and (oz) axis then: -

−−= cos)( soz PP (1)

Then, as known that from the geometrical shape of the elements xS =cos Where, the value of

xoz PP )( −−= (2)

Basically, this is trough for only element that occurs in upper surface of the aerofoil wing section model. So, for

the element at the lower surface of the airfoil wing section model, it will become:

xoz SPPS )( −−= (3)

And, by integrated this through the (x) at a limit of

x = 0 to x = c, then, the full integration of the value z will be the total value of the force z[N], hence,

−+−−=

c

o

c

o dxPPdxPPZ00

)()( (4)

For the upper and lower airfoil wing section model surface, respectively, and by using subscripts ( )u and ( )l this

leads to:

dxPPPPZ

c

louo −−−=0

)()( (5)

These calculations will give the explanation of the variation of pressure (P) in all over and along the chord of the

airfoil wing section model, also, it is possible to estimate or calculate the lift. Even if the estimated pressure that

holds inside the aerofoil wing section model could be assumed to be (Po), however, the real value will be quite

irrelevant next step. In the next step, it will show the reason why the value of (Po) was really to be taken. The 4th

equation is easily will represent by a coefficient of Cz. which is defined by: -

Sv

zCz 22/1

=

And by consideration a unit span, then, the chord (c)[mm] is equal to the area (s)[mm2].

dxPPPPCv

Cv

zC

c

louo

z

−−−−

=

=

0

2

2

)()(2/1

1

2/1

(6)

Then by dividing 22/1 v into the inside of the square bracket will leads to, knowing that

( )

=

c

xddx

c

1 Gives: ( )

−−= c

xdCCC plpuz

1

0

(7)

Since, P

o Cv

PP=

−22/1

by definition.

So, by assuming a same step to be use in order to figure the formula .sinsin)( zssox PP =−=

which will lead finally to


20

=

cz

cz

xc

zdCpC

/

/

2

1

. (8)

Basically, depending on plane area of the aerofoil model section is related to this coefficient. Also, its suitable

to calculate the pinching moment from the points of distribution of pressure and/or for the effortlessly and simple

of calculation, only about the origin of the (ox) and (oz) axes, it will be found. So, respectively.

dxPPPP louoz )()( −−−−=

Hence, the contribution to be applied to the pitching-moment arrived according to this slice part of the z…. the

force is as follow: -

dxxPPPP louoM .)()( −−−= (9)

Hence, total pitching-moment that happened according to the force (Z) is: -

−=

−+=

c

c

luzM

c

xd

c

xCp

c

xd

c

xCpCpC

0

0

(10)

Since, 222 2/1.2/1 Cv

M

cSv

MCM

=

=

where, in this case S = c, also,

In the sequence, the value MC contribute through the x…. force and it also might be found as: -

= c

zd

c

zCpC

cz

cz

xM

2

1

(11)

Then, the integration above is normally performed by graphic drown. The knowing that the force coefficient (Cx)

and (Cz) are perpendicular and parallel to the estimated chord line, where the more aerodynamic characteristics

usual coefficient (CL) and (CD) were usually referred to the direction of air. And the conversion from one pair to

anther might be also performed by the reference to (Figure-2), in which (CR), and the coefficient of the resultant

of both (Cx, Cz, CL and CD) respectively, therefore, from (Figure-3) it will have found that: -

( )

sinsincoscos

cos

−=

+=

RR

RL

CC

CC

And: xRzR CCandCC == sincos Where

sincos xzL CCC −= (12)

Similarly,

( )

cossin

sin

xz

RD

CC

CC

+=

+= (13)

Since,

−=

q

PPC p (14)

Where 2v

2

1q = (15)


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Then:

−=

q

PPC u

puAlso

−=

q

PPC L

pL (16)

Where, the pressure (P) is usually represent experimentally by (h) [1].

Figure 2. Represent an Aerofoil section model in the position of the fluid stream, at a selected speed of (V).

Figure 3. The determinant resultant of forces acting over the Aerofoil wing section model.

SUBSONIC WIND TUNNEL

In this study, NACA 0015 wing section model was used with internal pipe tubes installed as shown as in Figure

(4), while the experimental subsonic open channel type wind-tunnel was used in the current research of the suitable

type of an open circuit, that use a cross sectional visible working area of (30 cm x 30 cm), as shown and represented

simply in Figures (5-6). The internal wind speed of maximum of (36 m/sec.) is simply achievable by motor moves,

which allow the experiments in many speeds of airflow stream velocity and, hence, a subsonic aerodynamics is

said to be performed at acceptable values. The tunnel has assumed as frictionless wall with glass viewed test

section of (20 cm x 20 cm) and thickness of 5 mm. Figures (7-8) simply represent the sketches of image of the

wind-tunnel and the visible cross section working, respectively.

M

R

C

V


22

Figure 4. The aerofoil wing-section model of (NACA 0015) with the pressure vessel tubes of stainless-steel

fixed inside and through it

Figure 5. The open test section subsonic wind tunnel designed with digital flowmeter.

With respect to this part of the working model section, at the diffuser which essentially will leads to airstream flow

from a selective fan drive by a (6 kW and motor type of three phases A.C one). The airstream flow is basically

assumed to be controlled by a changeable valve type before the exit position in to the atmospheric throughout the

exit duct. Basically, the air/fluid were usually entering the wind-tunnel through a well design determined diverging

converging duct. The full visibility of the working section model is garneted all over the flow field, and the model

section is assumed to be supported only from one side of the wall and considered as cantilever. For the point at the

downstream, there is a selective pressure measurement device type digital pitot static tube, Figure(7) represent the

selective devise of pitot tube that been used in this work with the wind-tunnel working section open channel type,

where, the selective devise used of pitot-tube is associated through the wind-tunnel at the upper position. Then the

thickness of the boundary-layer was found by using the empirical formulas below for both the laminar as well as

the turbulent boundary-layer that will represented respectively as the following: -

Figure 6. Multiple models designed for the wind tunnel


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Figure 7. The correction factor curve for pitot and the digital pitot-tube used

x

lam

x

Re

64.4. = Represent the laminar flow on the airfoil surface position (17)

5..

Re

37.0

x

turb = Represent the turbulent flow on the airfoil surface position (18)

The change in the boundary layer from laminar into a turbulent level was considers as a precipitation change at

Rex = 500000, in a way that the thickness of the turbulent boundary layer was found to be approximately equal to

26 mm at the test section exit position which is simply assumed to be acceptable and does not affecting a lot at the

test section position. Even so, the turbulence flow level at the working section was simply found to be extremely

too acceptable [1].

WIND-TUNNEL CORRECTIONS FOR BOUNDARY LAYER

For aerofoil section model, the boundary conditions in any wind-tunnel at the time of test is generally not the same

as the condition that the aerofoil section model is in a free air. Basically, because of the walls on the airfoil section

model thickness, the wake blocking and finally the wake that subjected to the solid, which are usually can be

assumed to be neglected with model type of an open channel test section, since the airflow stream is then assumed

to be free to flow in a normally, we could have assumed that: -

areationTest

areafrontalModelCt

sec2

1=

Also, it was found that an optimum ratio of the airfoil model section frontal area to the cross-sectional testing area

of the aerofoil model section was about 7.5% need to be applied, unless for the condition that there were some

kind of mistakes of a several percent's that can be really accepted.

In this work, the aerofoil wing section selected model of a frontal area to the test cross sectional area was

approximately about almost 0.85%. That will means blocking the mistakes are relatively so small that irrelevant.

Basically, relevant to aerofoil wing section models of a large ratio that to be more than about 7.5% this will lead

to the conclude of the solid-blocking equation (18), also, the wake-blocking of equation (19) corrections that were


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represented and calculated respectively as the following empirical eq.: - 2

3

msb A/)mv(K=

(19)

Where,

Km : Airfoil wing section model shape factor = 0.74 for the selected model

mv : Airfoil wing section model volume = 0.7.t.C.b[mm3]

A : Airfoil wing section model test sectional area = 93025mm2 for the selected wind tunnel.

And,

dun

t

wb Ch

C.

2

1= (20)

Where,

ht : Heights of the selected wind tunnel test section used.

Cdun : The uncorrected drag coefficient estimated.

Then, the overall correction coefficient will be as following: -

sbwb += (21)

Also, the Rex correction arrive from this equation is as below:

)1(ReRe +=unxx (22)

Then, the correction of lift coefficient will be:

)21( −−=unLL CC (23)

Where:

22

48

=

ht

C (24)

Also, correction in the drag coefficient will be as following:

)231( wbsbundd CC −−= (25)

Finally, the correction in the angle of attack will be:

+

+=

un

munun CCL4

142

3.57

(26)

Where, un is representing reads that uncorrected [5].

THE WIND TUNNEL SECTION MODELING

By the application of CFD (Computational Fluid Dynamics) and FEM (Finite Element Method) Figure-6, a

numerical mathematical approximation of the selected wind tunnel and airfoil wing section model of type NACA

0015 was mathematically as well as experimentally built and the airfoil wing section model was summitry type as

well-known of this NACA series type, i.e. the lower surface is equal exactly to the upper section. Both data for

this wind tunnel and airfoil wing section model were driven from lists of the NACA's models that is previously

available [9], [10]. The specification of the airfoil wing section model was as following:

a. The cord C = 150 mm,

b. The thickness ratio of about 15%, and,


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c. The section length (b = 305 mm).

Figure 8. CFD FEM Design Model of the Wind Tunnel

CORRECTIONS IN PITOT-STATIC TUBE

A standard Pressure vessel of pitot-static tube with standard elliptical head shape type, (N.P.L Standard) was used

in this study, in order to estimate the arrived airspeed of the flow inside the selected wind-tunnel that been used.

Taking its diameter of about 5mm as well as about 20 cm long. And, the empirical equation for correction the

pitot-static tube as [12]:

2

1

...2

1

.)(

=−=

=

uEPP

gHP

statictotal

water

(27)

++= oEE (28)

Where,

E : Represent the factor of correction applied.

Eo: A Distance effects of the inlet static pressure (Eo=0.9976).

ω : The effect of viscosity, hence, for the fully developed flow the value of ω = 0.

Ω : The distance effect estimated from the wind-tunnel wall to the tube, and it might be estimated directly from

the schematic curve as in Figure (7).

MULTI TUBES MANOMETERS

In this work, s standard manometer tubes were selected to estimate pressure of the airstream airflow arrived Figure

(5), and the liquid to be selected for this type of manometers tubes was simply water, that because of testing were

in subsonic criteria only. So, no need for mercury. A rubber snout was related to the all stainless-steel tubes

manometer of the aerofoil wing section model.

RESULTS

CFD package, using the fundamentals of FEM approximation was used to estimate and calculate the simulation

of the wind tunnel by Ansys work bench. And the results of the pressure distribution are as follow:

1. Angle of attack (α)=5. See Figure (9)


26

Figure 9. CFD 2D Model of Wind tunnel with NACA 0015 and α = 5o

2. Angle of attack (α)=10. See Figure (10)

Figure 10. CFD 2D Model of Wind tunnel with NACA 0015 and α = 10o

3. angle of attack (α)= -10. See Figure (11)

Figure 11. CFD 2D Model of Wind tunnel with NACA 0015 and α = -10o


27

Figure 12. Pressure distribution and Aerodynamic characteristics coefficients for NACA 0015 with respect to

effective angle of attack αeff = 5o

Figure 13. Pressure distribution over the upper surface for NACA 0015 with critical ratio effective of about 75%

from the chord and angle of attack αeff = 5o

DISCUSSION AND CONCLUSION

From this theoretical -experimental work, it could be concluded: -

1. The aerodynamic characteristics (Cpu, CL, CD) is enhanced with increasing and/or decreasing of the angle of

attack angle, and the stalling angle for this NACA wing section was at αs = -10o and 10o (see Figures (9 ~ 13)).

This is due to the flow separation over the aerofoil wing section model.

2. For the same reason above, the coefficient of moment Cm and the upper coefficient of pressure CpL, were

decreased with increasing of the angle of attack angle, and the effective angle of attack for this NACA was at αeff

=5o (see Figure (12)).


28

3. The critical ratio holds at upper position for the selected aerofoil wing section model of (NACA 0015) was

about (75%) from the chord, Figure (13), Illustrates the Stress, Strain and displacement distribution for the NACA

airfoil section models and it shows how the critical ratio over the upper position with effective angle of attack

equal to 5o.

4. The boundary layer separation of the airfoil wing model section at both upper and lower surfaces of (NACA

0015) starts with the decreasing or increasing of the angles of attack ratio and it was found that the value effecting

over the upper position of the aerofoil wing section model surface were little high according to its values that holds

lower one, even if the wing section is symmetric, which is the reason why most of controlling ways for airstream

separation and boundary layer designed where to be applied at the upper positions of wings, also, may be due to

less boundary layer separated effects occurs at the lower position of the wings, most of the modern airplanes

recently avoided putting the engines at their positions.

5. Figure (13), represent the effect of normal flow applied all over the surface of the aerofoil wing section model

concentrating on the weak position of the aerofoil wing section model, and the pressure distribution over the upper

surface for NACA 0015 with the critical ratio effective of about 75% from the chord and angle of attack αeff = 5o.

6. An investigating for the effects of changing aerofoil wing section model with or without controlling surface is

highly recommended.

7. Also, if possible, we strongly recommended to investigate the parameters of basic aerodynamic characteristics

for this airfoil model section for a transonic speed with swept back angles ratio changing.

REFERENCES

[1] AA Shandookh, et.al, “Experimental Investigation of Aerodynamic Characteristics of NACA 23015 under

different angles of attack and Comparison with Available Package” Engineering and Technology Journal, Vol.27,

Issue.2, PP 215-234, 2009.

[2] Pope & John, J. H., “Low Speed Wind Tunnel”, John Wiley & Son., New York / London. 2001

[3] E. L. Houghton P. W. Carpenter Steven Collicott Daniel Valentine, “Aerodynamics for Engineering and

Students”, Butterworth-Heinemann. 2016.

[4] C. Kermode et.al., “Mechanics of Flight”, Pearson Education Limited / Harlow, United Kingdom. 2012.

[5] Aerodynamic Lab. Technical Notes, Mechanical Department, University of Technology. Iraq 2017.

[6] Chue S. H., “Pressure Probes for Fluid Measurement”, Progress in Aerospace Sciences, Vol.16, No.2, pp.

147 – 223, 1975.

[7] Charles E. Dole, et.al. “Flight theory and Aerodynamics”, John Wiley .2016.

[8] Alan Pope, “Basic Wing & Airfoil theory”, Dover Publications Inc. New York, United States. 2009.

[9] H. Abbott & A. E. Van Danhof, “Theory of wing Section”, McGraw-Hill, NY. 1959.

[10] D. John Anderson, JR., “An Introduction to Flight”, 7th Edition. McGraw-Hill, 2011.

[11] J. John Bertin & L. Michael Smith, “Aerodynamics for Engineering”, Prentice Hall. 1989.

[12] N Kadir. “Modification of Airfoil Aerodynamic Properties by using Internal Aquatic Excitations Technique”

MSc. Theses, Mechanical Department, University of Technology. Iraq, 2005.

[13] L. J. Clancy, “Aerodynamics”, 9th Edition .2006.

[14] Markel, C. E., “A Theory of Blockage Effects of Buff Bodies and Stalled Wings in A Closed Wind Tunnel”.

ARC R & M 3400,1965.

[15] Royal air force – UK, “The Principle of Flight”, Vol. A, Sec.3, Ch.2,1992

[16] Royal Aeronautical Society “Journal of the Royal Aeronautical Society”, Vol.64, No.596, P.449-464,

Aug.1960 & Vol.61, No.553, P.37-42, Jan.1957

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