Hello everyone,

I have a few questions about the procedure of using Airfoilprep and some of the inputs.

I will state the procedure that i use and please correct me if i am wrong:

1- Since i have no access to experimental results for AOA, Cl and Cd i use Xfoil to calculate alpha Vs Cl and alpha Vs Cd for a range of AOA sat (from -3 to 23) .

2- I am using only one Re so i will skip interpolate worksheet.

3- I am only using only one airfoil so i will skip blended airfoil worksheet.

4- 3D stall worksheet:

I have checked the Selig Du equations below:

Where

r/R and Chord: Do i have to enter r/R at each station, for example if i divide the blade to 10 elements do i have to enter r/R 10 times and

make airfoil .dat file for each station, taking in consideration that the airfoil is the same at all sections?

Rotor RPM:The same question here: since the RPM change which RPM should i use?

Min and Max alpha : Is this values related to the range that i choose at the beginning?

Alpha end: How can i choose this value?

Now i just copied the AOA, Cl and Cd into the worksheet and calculated 3D table.

5- TableExtrap worksheet:

Aspect ratio: How can i calculate this value if the blade is twisted and which chord should i use?

I know that it said: This value (CD max) is the one we’ll use. The Aspect Ratio value is provided for reference only, but i see it in the

above equations!!

Now i copied alpha, Cl and Cd to the table and update table.

6- DynStall:

Min and Max alpha: The same question as in 3D stall worksheet.

Stall angle of attack: how can i get this value, did i get it from Cl alpha curve?

Cn at stall value for negative angle of attack: How can i get this value and which parameters control it?

Is it an important parameter because i notice that when i change it nothing happened neither to the curve nor

to the values?

Now update all.

7-Editor format:

I know it some kind of long but i hope this will become reference for others.

Thanks in advance.