Pressure Distribution Over an Airfoil

8/10/2019 Pressure Distribution Over an Airfoil

1/11

PRESSURE DISTRIBUTION OVER AN AIRFOIL

ME323: FLUID MECHANICS LABORATORY

Submitted By: Yousif Alromaithi

Submitted To: Joe

Submitted: 11/12/2014


2/11

Press. Distribution Over An Airfoil

2

ABSTRACT

The experiment is to be performed to determine the pressure distribution over an airfoil and use

it to determine the Coefficient of Pressure, Lift force, pressure drag, the lift and drag coefficients

CL and CD. A Clark Y-14 Airfoil and a wind tunnel were used in performing this experiment.

The difference in pressure over an airfoil causes a lifting force to it. The flow above the airfoil

becomes accelerated when it hits the airfoil and by applying the Bernoullis equation, we see that

the pressure decreases. While the flow above the airfoil is accelerated, the flow below it

decelerates therefore increasing the pressure and causing a lift to the airfoil due to the difference

of pressure.

The pressure distribution was determined at different angle of attacks. The pressure were

recorded using sensing holes on the airfoil. The coefficients of drag and lift were calculated

using the pressure recorded. The results were plotted for analysis.


3/11


3

DATA, THEORY AND RESULTS

Air Velocity: 60mph or 27 m/sLength of Airfoil: 11.375in or 0.288925m

Average Static: 0.45kPa or 0.06526psi

Static

Number

Percentage of

Chord

Pressure Data

(kPa)

0 0 0.37 0.37 0.21 -0.03

1 5 -0.08 -0.28 -0.51 -0.68

2 10 -0.21 -0.38 -0.57 -0.69

3 20 -0.29 -0.41 -0.54 -0.6

4 30 -0.26 -0.34 -0.45 -0.53

5 40 -0.24 -0.3 -0.39 -0.35

6 50 -0.21 -0.27 -0.29 -0.3

7 60 -0.18 -0.23 -0.22 -0.23

8 70 -0.17 -0.12 -0.15 -0.15

9 80 -0.03 -0.05 -0.07 -0.07

10 5 -0.22 -0.05 0.1 0.19

11 10 -0.14 -0.03 0.08 0.15

12 20 -0.06 0.03 0.1 0.13

13 30 -0.01 0.04 0.09 0.11

14 40 0 0.04 0.08 0.1

15 50 0.01 0.04 0.07 0.09

16 60 0.01 0.04 0.06 0.08

17 70 0.01 0.04 0.06 0.06


4/11


4

Static Ring 0.45kPa 0.44kPa 0.46kPa 0.45kPa

Atop(in )

Locations 09[x*L]

Top Forces(lb)

[Pressure*Area]Convert P to psi

ABottom(in )

Locations 10-17[x*L]

Bottom Forces(lb)

[Pressure*Area]Convert P to psi

10.053664

- -

1.99 -0.08082 - -

2.99 -0.16479 - -

3.98 -0.23667 - -

3.98 -0.19627 - -

3.98 -0.17318 - -

3.98 -0.15586 - -

3.98 -0.13277 - -

3.98 -0.06927 - -

3.98-0.02886

-

-

- - 1

-0.00725

- - 1.99 -0.00866

- - 3.98 0.017318

- - 3.98 0.02309

- - 3.980.02309

- - 3.98 0.02309

- - 3.98 0.02309


5/11


5

- - 3.98 0.02309

The charts are results for the zero degree angle of attack.

AFront(in )

Locations 0-5 &

10-13[y*L]

Front Forces (lb)

[Pressure*Area]

Convert P to psi

ABack(in )

Locations 5-9 &

13-17[y*L]

Back Forces (lb)

[Pressure*Area]

Convert P to psi

Cp

1.070.05742044

- -

0.822222

2.49 -0.10112031 - - -0.62222

1.07 -0.05897234 - - -0.84444

1.07 -0.06362806 - - -0.91111

0.53 -0.0261358 - - -0.75556

-0.49 0.02132055 - - -0.66667

- - -0.460.01801369 -0.6

- - -0.69 0.02301749 -0.51111

- - -0.9 0.01566408 -0.26667

- - -0.43 0.00311831 -0.11111

0.78-0.00565647

- - -0.11111

0.11 -0.00047862 - - -0.06667

0.070.00030458

- -0.066667

0.00 0 - - 0.088889

- - 0.180.00104427 0.088889


6/11


6

- - 0.53 0.0030748 0.088889

- - 0.36 0.00208854 0.088889

- - 0.36 0.002088540.088889

Solving for the Normal Force and the Axial Force:

Solving for the Coefficient of Drag and the Coefficient of Lift for zero degree angle of attack

()

()

For 3,-3, and 5 degree angles, computed separately.


7/11


7

()

()

()

For plotting the CPvs x

Bottom x Top

0.822222 0.00

0.822222

-0.11111 0.18 -0.62222

-0.06667 0.35 -0.84444

0.066667 0.7 -0.91111

0.088889 1.05 -0.75556

0.088889 1.4 -0.66667

0.088889 1.75 -0.6

0.088889 2.1 -0.51111

0.088889 2.45 -0.26667

2.8 -0.11111


8/11


8

Plot of Cp vs x

Cp vs Y

y Cp

-0.14 -0.11111

-0.13 -0.06667

-0.13 0.066667

-0.13 0.088889

-0.13 0.088889

-0.09 0.088889

-0.06 0.088889

-0.06 0.088889

-1

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

0 0.5 1 1.5 2 2.5 3Cp

x(in)

Top

Bottom


9/11


9

0

0.822222

0.19 -0.62222

0.19

-0.84444

0.23 -0.91111

0.25 -0.75556

0.29 -0.66667

0.31 -0.6

0.32-0.51111

0.31 -0.26667

0.38 -0.11111

ANALYSIS AND DISCUSSION

The experiment aims to observe and calculate the pressure distribution over an airfoil. The

pressure around the Clark Y-14 airfoil were recorded for different angles of attack, 0, -3, 3, and

5. The airfoil was placed in a wind tunnel with air velocity of 60 mph. The Coefficient of drag

and Coefficient of lift were determined for the airfoil with an angle of attack of zero degrees.They were determine by calculating first the lift force and drag forces acting on the airfoil.

The results showed difference between the top and bottom faces of the airfoil and also the back

and front faces of it. The force acting on top is negative while the force on the bottom is positive.


10/11


10

The coefficients of pressure for the top face are almost all negative compared to the bottom

which all have positive values.

CONCLUSION

The experiments objective is to show the pressure distribution over a Clark Y-14 airfoil

subjected to air stream. The pressure values were recorded from 0 to 70% of the chord length.

The differences of the pressures around the airfoil causes it to have a normal or lift force and an

axial or drag force. The differences were caused by its shaped which alters the velocities of the

airs stream.

The lift also increased as the angle of attack increases. I believe this is the case only if the angles

are small. In larger angles, the lift is not directly proportional to the angle of attack. The

coefficient of pressure were also calculated and plotted with the values of x. The plot of Cp is

almost the same as the sample plot.

PRACTICAL APPLICATION

Anairfoil is a streamlined shape that is capable of generating significantly more lift than drag. A

flat plate can generate lift, but not as much as a streamlined airfoil, and with somewhat higher

drag. Airfoils have lots of applications in the field of engineering. From propellers, fans,

turbines, and specifically for wings of planes. Different types of airfoils are used for different

uses of it. Each has different shape and effect as fluid flows over them.

The study of airfoils are very important since they are used in planes that can endanger the life of

the passengers if not properly used. Airfoils are very important to aerodynamics and should be

studied.
http://en.wikipedia.org/wiki/Airfoilhttp://en.wikipedia.org/wiki/Airfoil


11/11


11

REFERENCES

1. http://www.grc.nasa.gov/WWW/k-12/airplane/incline.html

2. http://hyperphysics.phy-astr.gsu.edu/hbase/fluids/angatt.html

3. http://web.mit.edu/2.972/www/reports/airfoil/airfoil.html

4.

http://hyperphysics.phy-astr.gsu.edu/hbase/fluids/airfoil.html5. http://www.grc.nasa.gov/WWW/k-12/airplane/lifteq.html

6. http://www.aviation-history.com/theory/airfoil.htm
http://www.grc.nasa.gov/WWW/k-12/airplane/incline.htmlhttp://hyperphysics.phy-astr.gsu.edu/hbase/fluids/angatt.htmlhttp://web.mit.edu/2.972/www/reports/airfoil/airfoil.htmlhttp://hyperphysics.phy-astr.gsu.edu/hbase/fluids/airfoil.htmlhttp://www.grc.nasa.gov/WWW/k-12/airplane/lifteq.htmlhttp://www.grc.nasa.gov/WWW/k-12/airplane/lifteq.htmlhttp://hyperphysics.phy-astr.gsu.edu/hbase/fluids/airfoil.htmlhttp://web.mit.edu/2.972/www/reports/airfoil/airfoil.htmlhttp://hyperphysics.phy-astr.gsu.edu/hbase/fluids/angatt.htmlhttp://www.grc.nasa.gov/WWW/k-12/airplane/incline.html

Pressure Distribution Over an Airfoil

Documents

Transcript of Pressure Distribution Over an Airfoil