8/6/2019 Term Project 2

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FLOW OVER A NACA AIRFOIL

Part 1: Creation of geometry

http://www.ppart.de/aerodynamics/profiles/NACA4.html

Visit the site given above to calculate the shape of the NACA profile you are assigned. (The fileParametersPROJECT2.pdf gives the name of your NACA profile)

1) Create a WORD document by copy and paste.

2) Change points to commas and eliminate some points (involving Exponential terms) in the WORD file.3) Copy the X,Y data of the WORD file to the C column of an EXCEL file. Separate this column into two columns

(C and D) bu using: Data, Convert Text into Columns and Space in three steps). Enter ones to Column A(Part

Number is one). Enter (1,2,3,.. Number of data points) in Column B. Finally add zeros (z coordinate of the

profile) to Column E.

4) Create a data (.dat) file by using copy and paste. This file can be read by ANSYS Fluent. An example data file

(NACA4312.dat) is attached in Class notes of April 11th.5) Copy your data file to a removable memory device and bring it to the lab on APRIL 25th.

Part 2: Solution and analysis

Solve the flow over your NACA airfoil for the following 3 cases:

A) = 0 degrees (Angle of attack), Laminar flow

B) = 10 degrees (Angle of attack), Laminar flow

C) = 10 degrees (Angle of attack), Turbulent flow

For all cases: Use Double precision. For iterative convergence, monitor lift coefficient CLand make sure that CL is approximately constant (within plus minus 0.01) when iteration is stopped.

For Cases A and B , use U=1 m/s, =0.001 kg/m.s, =1 kg/m^3 (Re=1000)

For Case C, use U=1 m/s, =0.000001 kg/m.s, =1 kg/m^3 (Re=1000000)

Use the following velocity inlet conditions:

For Case A: UX =1 m/s, UY =0 m/s

For Cases B and C: UX = U (Cos 10)= 0.985 m/s, UY = U (Sin 10) = 0.174 m/s

Reference Values: Area=1 m^2, =1 kg/m^3, =0.001 kg/m.s (for cases A and B)

=0.000001 kg/m.s (for case C). Note that reference values effect nondimensional

quantities (such as Lift coefficient (CL), Drag coefficient (CD), Pressure coefficient (CP, ) only)

Save 4 pictures (Contour plots of pressure and UX and also plots of Residuals and Lift History)

for all three cases. Also write down the values of Lift (L), Drag (D), Lift coefficient (CL) and Drag

coefficient (CD ) for all three cases.

Compare your results of three cases studied in terms of:

1)Contour plots of pressure cofficient and streamwise (x) velocity (Present plots for all three cases on the same

page of your report. Field of view must focus around the airfoil. Present also plots for the Residuals and Lift

History)

2) Values of Lift (L), Drag (D), Lift coefficient (CL), Drag coefficient (CD ) presented as a table for all cases. (Note

that when =1 kg/m^3, U=1 m/s and Area=1 m^2, CL=2L, CD=2D )

http://www.ppart.de/aerodynamics/profiles/NACA4.htmlhttp://www.ppart.de/aerodynamics/profiles/NACA4.htmlhttp://www.ppart.de/aerodynamics/profiles/NACA4.html