aerofoil assignment.pdf

KINGSTON UNIVERSITY-K0827514

1. INTRODUCTION:

The birth of wind tunnels was even before the Wright Brother’s success in 1903. It

was first designed and operated in the year 1871 by Frank H. Wenham. Since then

great advancement and understanding took place in airfoil profiles and designs.

( www.grc.nasa.gov, 2009)

Wind tunnels are primarily used to study aerodynamic effects on objects under test.

In aeronautical engineering, they provide facility to test proposed designed models of

aircrafts and parts, duplicating the various aerodynamic effects on a full-scale

aircraft. This enables designers to improve performance and features of the

designed model before stepping onto a full-scale production. Thereby cutting down

unnecessary costs involved, they also allow the possibility of comparison of different

aerofoil shapes and determine the behavior of air around the airfoil as well. Purpose

of wind tunnels is to create a low-turbulent, high-speed airflow through the test

section in order to obtain precise values of lift and drag data that later used for

analysis. (www.fi.edu, 1998)

The experiment was carried out on a Pitsco Aitech 40, which is an open circuit type

wind tunnel. The tests were done on six different aerofoil in order to achieve a sound

knowledge of aerodynamic effects on the test piece. Some of the objectives of this

experiment also include finding of stall angles and the angle of attack where the wing

is most efficient. Finally, above all it also highlights recommendation on improving

wind tunnel construction in order to obtain better results.( www.grc.nasa.gov, 2009)

http://www.grc.nasa.gov/

http://www.fi.edu/

http://www.grc.nasa.gov/


2. TEST METHODS:

The experiment is done with a view of finding data such as Lift, drag forces and

particular stall angles for each aerofoil. During the experiment, six different aerofoils

are used. The experiment involves 3 tasks to be completed. First, it requires the

experimenter to find out the stall angles for each aerofoil using the graph computed

from wind tunnel software. Then use values of the data plotted by the computer to

calculate lift (CL) and drag (CD) coefficients by the use of formula given below:

CL =

⁄

CD =

⁄

(www.centennialofflight.gov, n/k)

Finally, the angle at which the wing is most efficient is calculated through lift/drag

ratio.

3. EXPERIMENT IMPLEMENTATION:

To begin with, the class was broken down to six groups of four. Then each group

was assigned to carry out test on one of the six foils in the wind tunnel. This involves

the experimenter to securely mount the aerofoil at 00 to the horizontal. Once verified

through visual inspection, the experimenter then proceed on to configure the

computer to test angles from 00 to 450 . Its important make sure those readings are

not taken immediately after the tunnel is switched on. But allow the tunnel run free

for few seconds in order to stabilize the airflow over the wing. Then after satisfactory

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amount of time the readings are taken and by taking the mean of a number of

measurements for each aerofoil reduces the random error caused. It also important

to make a note that environment characteristics where assumed to be at sea-level

(e.g density,pressure) while the experiment was carried out. After every satisfactory

test, the graph is saved onto the computer which is later printed-out for analysis.

Next, the dimensions of the six models are recorded to determine area, chord length,

etc. Lastly, above procedure is repeated for all six aerofoil separately.

Pic 1:Wind Tunnel

By :Sammy Ritoch


4. TEST LIMITATION:

It’s important to take into account the tunnel by itself has limitation that restricts the

experimenter from performing fully defined operation. This includes availability of

space which limits the size of aerofoil that could be tested. Moreover, to have

controlled laminar flow over the wing could be difficult owing to course surface finish

and surface contamination through dust and dirt. These create turbulent airflow

within the tunnel. Another factor is the Tunnel vibration generated by running

propeller. This causes buffeting of the aerofoil contributing an error factor to results.

One other factor is direction of relative airflow at which it strikes the aerofoil. The

oncoming airflow hits the aerofoil at an angle due to converging design of the intake

ducts and thus giving negative lifts for first few angles resulting off-set of the graph

origin(by 4° approx). In addition to above, mounting of aerofoil 0° to horizontal was

challenging task and was confirmed merely through a visual inspection.

( www.fortus.com, 2010)

5. RESULTS

Refer to Appendix A and B for values obtained through the experiment.

The values were produced in terms of force (Oz) which requires the following

formulas given below to find CL and CD coefficients

CL =

⁄

CD =

⁄

Where V=velocity, = density of air, S= surface area, D=drag,L= lift

http://www.fortus.com/


5.1. AEROFOIL 1:

Pic by: West. J Pic 2

Above graph 5.1 shows Coefficient of Lift/Drag characteristics of aerofoil 1

This particular aerofoil shows maximum lift is obtained approximately 8° and starts to

stall soon after. The lift of this aerofoil rises almost steadily up to stall angle while the

drag shows no adverse change with stall angle.

0.00

0.01

0.02

0.03

0.04

0.05

0.06

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

Angle of Attack

Aerofoil 1

CL

CD


5.2. AEROFOIL 2



This aerofoil generates lift until about 12° and then looses lift dramatically. While the

drag curve increases at slow rate with increasing angle of attack.

-0.01

0.00

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

Angle of Attack

Aerofoil 2

CL

CD


5.3. AEROFOIL 3



It is obvious from the above graph this aerofoil is quiet unique. It has two points of

noticeable stall. It begins its 1st stall around 5° and then catches lift again at about

6.5°. Final stall occurs between 10-11° and thereafter looses lift significantly. In

addition, adverse negative lift is observed during initial angles of attack.

-0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

Angle of Attack

Aerofoil 3

CL

CD


5.4. AEROFOIL 4



It is visible from above graph it has a fairly higher stall angle of 17 odd degrees and

also produces greater lift coefficient of 0.104. Therefore the graph has an extended

scale to accommodate higher values. Drag coefficient remains zero until 4° and kicks

in between 4-5° and rises steadily with minor fluctuation.

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0 2 4 6 8 10 12 14 16 18 20

coe

iffi

en

t

Angle of Attack

Aerofoil 4

CL

CD


5.5. AEROFOIL 5



Aerofoil 5 shows a steady increase in lift coefficient up to stall angle of 13°. However

there is no massive loss in lift but a gradual decline. On the other hand drag remains

fairly straight and increases linearly with increasing angle of attack.

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

Angle of Attack

Aerofoil 5

CL

CD


5.6. AEROFOIL 6



The fact this aerofoil produce the most lift compared to all other aerofoil is proven

from above the graph. This aerofoil generates a maximum lift coefficient of 0.116 and

starts to stall at high angles of attack in close proximity to 14°.

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0 2 4 6 8 10 12 14 16 18

Co

eff

icie

nt

Angle of Attack

Aerofoil 6

CL

CD


5.7. LIFT/DRAG RATIO- MEASUREMENT OF AERODYNAMIC

EFFICIENCY.

In order to determine the efficiency of aerofoil, the experimenter requires finding

CL/CD ratio by dividing the calculated figures of CL by CD. However, by equation and

theoretically Lift and drag forces are proportional CL and CD respectively and

knowing for the fact that conditions such as velocity of airflow, density and surface

area remains same throughout the experiment. We can determine the lift/drag ratio

by dividing the lift and drag forces directly by eliminating the constants. (See

Appendix C Efficiency state of a wing)

5.7.1. Aerofoil 1

Fig 5.7.1

Maximum efficiency lies at 2°

0

5

10

15

20

25

0 2 4 6 8 10 12 14 16 18

Lift

/Dra

g ra

tio

Angle of attack

Aerofoil 1


5.7.2. Aerofoil 2

Fig 5.7.2

Maximum efficiency lies at 7°

-5

0

5

10

15

20

0 2 4 6 8 10 12 14 16

Lift

/Dra

g ra

tio

Angle of attack

Aerofoil 2


5.7.3. Aerofoil 3

Fig 5.7.3

Maximum efficiency of the wing lies at 3°. The constant downhill slope is due to

generation of negative lift during the initial angles of attack.

-30

-20

-10

0

10

20

30

40

0 2 4 6 8 10 12 14 16 18Lift

/Dra

g ra

tio

Angle of attack

Aerofoil 3


5.7.4. Aerofoil 4

Fig 5.7.4

Maximum efficiency of this aerofoil lies at sharp 4°.

-10

0

10

20

30

40

50

60

70

0 2 4 6 8 10 12 14 16

Lift

/Dra

g ra

tio

Angle of attack

Aerofoil 4


5.7.5. Aerofoil 5

Fig 5.7.5

Maximum efficiency of the 5th aerofoil lies at 3°.

-5

0

5

10

15

20

25

30

35

0 2 4 6 8 10 12 14 16 18

lift/

Dra

g ra

tio

Angle of attack

Aerofoil 5


5.7.6. Aerofoil 6.

Fig 5.7.6

Maximum efficiency lies at high angle of 1°.

6. ANALYSIS

Aerofoil 1 and 2 belongs to symmetrical shape aerofoil. Therefore both produce zero

lift at zero angle of attack. However, close look on aerofoil 1 reveals it has slight

camber giving it a shape of a semi-symmetrical aerofoil. Hence it produces a slight

lift at zero angle of attack compared to aerofoil 2. When comparing the efficiency of

wings, aerofoil2 is less efficient than aerofoil 1(ref fig 5.7.1 and 5.7.2). This due to

fact that aerofoil 2 has greater thickness as compared to aerofoil 1, thereby creating

more drag when it rips through air than aerofoil1. Moreover use of aerofoil 1 has an

added benefit maintaining considerable amount of lift even after stalling as

0

2

4

6

8

10

12

14

16

0 2 4 6 8 10 12 14 16 18

Lift

/Dra

g R

atio

Angle of attack

Aerofoil 6


compared to aerofoil 2.(refer 5.1 and 5.2). Another drawback with the use of aerofoil

2, the high angle of attack at which wing being most efficient (7°) and low stall angle

(12°), thus limiting the amount of pitch it can attain during flight.

Moving to Aerofoil 3, the under-cambered design of the wing enhances the

production of lift even at zero angle of attack contrast to what is shown on the graph.

This due to experimental error, whereby the chord line of the aerofoil was mounted

facing below the horizontal axis. This error resulted in generation of negative lift at

zero angle of attack. The concave design on the lower surface of the aerofoil creates

a diverging path for the oncoming airflow and producing high-pressure underneath

the wing. Thus inducing more lift than usual. The design itself is a combination of 2

aerofoils giving 2 noticeable stall points (ref graph 5.3). (www.blackflight.com, n/k)

Taking the Graph 5.4 and 5.7.4 of aerofoil 4 into consideration we can deduce this

aerofoil produce the highest efficient of all other aerofoils and also maintains lift even

at higher angles of attack (17°). Moreover, this graph proves theoretical curve of

lift/drag ratio of the wing being most efficient at 4° of angle of attack. The reason for

its extraordinary efficiency and higher stall angles is the position of the maximum

camber of the wing which occurs at 50% of the chord line. This helps to maintain

laminar flow of air over greater distance across chord line. Laminar flow means less

drag and hence less energy is consumed making the wing more efficient during

cruise. By having the maximum camber near to middle of the wing also creates even

pressure distribution over wing surface. Therefore has a favourable pressure

gradient across the pressure recovery region (see the glossary) and thus having

more control over the onset of a transition point. These factors help to attain higher

angles of attack. Whereas aerofoil 5 is conventional type of aerofoil with camber

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situated at 25% of the chord line (ref fig 5.5). This results in an adverse pressure

gradient and a larger drag.( members.tripod.com, 2003)

From the analysis of graphs, it’s evident that aerofoil 6 produce the most lift. This

attributed to its highly cambered shape of underside of the wing. However, it’s least

efficient of all (ref 5.7.6).

7. CONCLUSION

It can be concluded by referring to appendix E the experiment proves the wings are

most efficient at about 4° as expected. The experiment also supported the fact that

Irrespective of the aerofoil shapes all wings begin to lose lift when it hits the stall

angle. It also highlighted the wing profile had an effect on the characteristics of

lift/drag ratios and stall angles of the wing. Lastly, the importance of using wind

tunnels for testing different aerofoil shapes to establish their behaviour when

subjected to airflows was satisfactorily appreciated and met with success.


8. RECOMMENDATION:

During the experiment implementation, number of factors affected to the accuracy of

results and performance of the wind tunnel. One such factor was the oncoming

airflow striking the aerofoil at an angle generating negative lift for the first few

degrees. This could be resolved by constructing a longer venturi tube, hence

keeping the test piece at further distance apart from the converging duct. This helps

the airflow to straighten-up by the time it reaches the aerofoil. In addition to this,

mounting additional honeycomb structure and wire mesh smoothing screen into the

converging duct would serve additional ways of achieving accurate and laminar

airflow in the test section. Another problem encountered during the experiment was

wobbly reading caused by aerofoil flutter. Replacing the softwood aerofoil with much

heavier and rigid aerofoil made from hardwood could overcome the flutter to serve

better results. One other reason for fluttering is improper positioning of centre of

gravity on the aerofoil. This is caused when centre of gravity is offset from hinge axis

of the aerofoil and causes the aerofoil to overshoot the equilibrium position due to

inertia. This is rectified by balancing the entire mass of the aerofoil at hinge axis.

Modification to test section can be made by making it more air tight so as to ensure

no suction takes place through window sealing. In the light of this conclusion, it

would be highly recommended to take the following measures stated above into

consideration for more better and accurate results in future. (www.fi.edu, n/k)

http://www.fi.edu/


9. REFERENCE:

Author n/k (1998). The Wind Tunnel Parts. NASA Observatorium. [Online]

Available at <http://www.fi.edu/flights/first/tunnelparts/index.html> [Accessed

on 01st August 2010]

Author n/k (2007). Air Pressure Distribution. Digital Textbook. [Online].

Available at <

http://www.desktop.aero/appliedaero/airfoils1/airfoilpressures.html>

[Accessed on 03rd August 2010]

Brendon J. (2009). Knowing the Aircraft. Fly Safe. [Online]. Available at < http://www.auf.asn.au/emergencies/aircraft.html>. [Accessed on 01st August

2010]

Ewing J. (n/k). Introduction to R/C Aircrafts. Black Flight Models. [Online]

Available at < http://www.blackflight.com/intro_rc/intro_rc_air.asp>. [Accessed

on 04th August 2010].

Ghods M. (2001). Wind Tunnel Testing. Theory of Wings.[Online] Available at

< http://members.tripod.com/m_ghods/frme2.pdf>. [Accessed on 05th August

2010] Talay, Theodere A. (not known). Subsonic Airflow Effects – The Two-

Dimensional Coefficients. [Online] Available at <

http://www.centennialofflight.gov/essay/Theories_of_Flight/Two_dimensional_

coef/TH14.htm> [Accessed on 30th July 2010]

William. R et al. (2009). Whirling Arms and The First Wind Tunnel. Wind

Tunnel of NASA. [Online] Available at < http://www.grc.nasa.gov/WWW/K-

12/WindTunnel/history.html>. [Accessed on 20th July 2010]

Wind tunnel(Pic 1) photo taken by Sammy Ritoch on 30th July 2010 at KLM

Technical college.

Aerofoil (Pic 2-7) photo taken by Joe West on 30th July 2010 at KLM

Technical College.

10. BIBLIOGRAPHY.

Author n/k (2009). Airfoil lnvestigation database. A.I.D.[Online] Available at <

http://airfoils.worldofkrauss.com:8888/web/help> [Accessed on 05th August

2010].

Author n/k (2010). Laminar Airfoil Theory. The Aviation History

Online.com.[Online] Available at < http://www.aviation-

history.com/theory/lam-flow.htm>. [Accessed on 05th August 2010]

http://www.fi.edu/flights/first/tunnelparts/index.html

http://www.desktop.aero/appliedaero/airfoils1/airfoilpressures.html

http://www.auf.asn.au/emergencies/aircraft.html

http://www.blackflight.com/intro_rc/intro_rc_air.asp

http://members.tripod.com/m_ghods/frme2.pdf

http://www.centennialofflight.gov/essay/Theories_of_Flight/Two_dimensional_coef/TH14.htm

http://www.centennialofflight.gov/essay/Theories_of_Flight/Two_dimensional_coef/TH14.htm

http://www.grc.nasa.gov/WWW/K-12/WindTunnel/history.html

http://www.grc.nasa.gov/WWW/K-12/WindTunnel/history.html

http://airfoils.worldofkrauss.com:8888/web/help

http://www.aviation-history.com/theory/lam-flow.htm

http://www.aviation-history.com/theory/lam-flow.htm


APPENDIX A: EXPERIMENTAL VALUES FOR CL AND CD

Aerofoil 1 AEROFOIL 2 AEROFOIL3 AEROFOIL 4 AEROFOIL 5

Angle of attack

CL CD CL CD CL CD CL CD CL CD

0 0.0051 0.0004 0.0005 -0.0002 -0.0067 -0.0002 -0.0112 0.0022 -0.0010 0.0000

1 0.0124 0.0007 0.0034 0.0002 -0.0020 -0.0005 -0.0045 -0.0022 0.0057 0.0000

2 0.0203 0.0009 0.0109 0.0009 0.0057 -0.0002 0.0915 0.0000 0.0121 0.0010

3 0.0266 0.0018 0.0164 0.0014 0.0158 0.0005 0.1339 0.0067 0.0220 0.0007

4 0.0322 0.0021 0.0193 0.0019 0.0309 0.0025 0.2767 0.0045 0.0275 0.0012

5 0.0335 0.0025 0.0268 0.0023 0.0361 0.0030 0.3547 0.0156 0.0322 0.0020

6 0.0429 0.0033 0.0298 0.0023 0.0237 0.0017 0.3525 0.0290 0.0334 0.0027

7 0.0505 0.0033 0.0405 0.0026 0.0275 0.0017 0.3882 0.0245 0.0393 0.0030

8 0.0561 0.0042 0.0423 0.0037 0.0356 0.0022 0.4730 0.0245 0.0507 0.0037

9 0.0531 0.0049 0.0528 0.0047 0.0401 0.0027 0.5667 0.0312 0.0564 0.0037

10 0.0529 0.0042 0.0480 0.0051 0.0673 0.0052 0.6381 0.0335 0.0646 0.0042

11 0.0452 0.0051 0.0546 0.0044 0.0646 0.0072 0.7296 0.0580 0.0695 0.0049

12 0.0501 0.0054 0.0582 0.0047 0.0492 0.0054 0.7162 0.0558 0.0685 0.0047

13 0.0494 0.0056 0.0243 0.0047 0.0304 0.0042 0.6983 0.0513 0.0769 0.0057

14 0.0424 0.0053 0.0212 0.0040 0.0289 0.0045 0.7073 0.0469 0.0730 0.0057

15 0.0463 0.0056 0.0207 0.0049 0.0235 0.0045 0.8121 0.0491 0.0591 0.0054

16 0.0482 0.0061 0.0189 0.0052 0.0228 0.0047 0.9192 0.0692 0.0552 0.0057

17 0.9080 0.0892

18 0.6740 0.0647

19 0.6870 0.0669

AEROFOIL 6

Angle of attack

CL CD

0 0.0423 0.0040

1 0.0509 0.0035

2 0.0621 0.0051

3 0.0659 0.0061

4 0.0673 0.0063

5 0.0612 0.0077

6 0.0752 0.0070

7 0.0846 0.0082

8 0.0885 0.0084

9 0.0841 0.0086

10 0.0834 0.0089

11 0.0920 0.0089

12 0.1047 0.0089

13 0.1096 0.0100

14 0.1140 0.0112

15 0.1065 0.0107

16 0.0930 0.0112

The following were the values of

variables used in the experiment:

Airflow Speed: 50mp/H

(22.22m/s)

Density Of Air: 1.225kg/M3

Surface Area: 0.1334 M2

Table 1


APPENDIX B: LIFT/DRAG RATIOS

Angle of attack

AEROFOIL 1 AEROFOIL 2 AEROFOIL 3

AEROFOIL 4 AEROFOIL 5

AEROFOIL 6

0 14 -2.00 27 -5.00 #DIV/0! 10

1 17 5.33 4 2.00 #DIV/0! 14

2 23 12.00 -22 #DIV/0! 12.25 12

3 15 12.00 32 20.00 29.67 10

4 15 10.63 12 62.00 22.20 10

5 13 11.80 12 22.71 16.25 7

6 12 13.10 13 12.15 12.27 10

7 15 16.18 15 15.82 13.25 10

8 13 11.63 16 19.27 13.67 10

9 10 11.60 14 18.14 15.20 9

10 12 9.59 12 19.07 15.35 9

11 8 12.63 9 12.58 14.05 10

12 9 12.80 9 12.84 14.58 11

13 8 5.35 7 13.61 13.52 10

14 8 5.47 6 15.10 12.83 10

15 8 4.33 5 16.55 10.86 9

16 7 3.61 4 13.29 9.70 8

17 10.18

18 10.41

19 10.27

Table 2


APPENDIX C: IDEAL THEORETICAL CURVE FOR EFFICIENCY

The graph to the right shows an ideal curve where the highest Lift-to-Drag

ratio most likely to occur. This usually occurs at

angle between 4° and 5°. At this point the wing

gains the maximum lift for minimum drag. Hence,

it is the optimum angle of attack for

cruise.(www.auf.asn.au, 2009)

APPENDIX D

Stall Characteristics;

Angle at which the rate of increase in lift starts to reduce.

Angle at which lift starts to decrease

Aerofoil 1 7.50 8.00

Aerofoil 2 11.60 120

Aerofoil 3 100 10.50

Aerofoil 4 16.20 170

Aerofoil 5 10.60 11.60

Aerofoil 6 13.40 14.40

Table 3

Brandon J. (2009). “Efficiency Graph”. Fly safe. [Online] Available at <

http://www.auf.asn.au/emergencies/aircraft.htmlhttp://www.auf.asn.au/emerg

encies/aircraft.html>. Accessed on 1st August 2010.

http://www.auf.asn.au/

http://www.auf.asn.au/emergencies/aircraft.htmlhttp:/www.auf.asn.au/emergencies/aircraft.html

http://www.auf.asn.au/emergencies/aircraft.htmlhttp:/www.auf.asn.au/emergencies/aircraft.html


APPENDIX E: AVERAGE POSITION OF HIGHEST EFFICIENCY

AEROFOIL NO

Aerofoil 1

Aerofoil 2

Aerofoil 3

Aerofoil 4

Aerofoil 5

Aerofoil 6

Average Angle of attack

Angle of Highest Efficiency

2° 7° 3° 4° 3° 1° 3.33°

Table 4

GLOSSARY

PRESSURE RECOVERY REGION:

This region of the pressure distribution is called the pressure recovery region.

The pressure increases from its minimum value to the value at the trailing

edge.( www.desktop.aero, 2007)

http://www.desktop.aero/

aerofoil assignment.pdf

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