[mohammad]

Dear all,

I have the results of a wind turbine's blade. I have created an airfoil-shape polyline (bye cross-sectioning the blade with a plane)
I can easily know the pressure distribution values on this polyline and also I can calulate the Cp.
Theoritically, If I want to use this Cp to calulate the "Lift" and "Pressure Drag" I need to calculate the integral of Cp.

However, the CFX-Post itself has a command " lengthInt". As read from the command description, it calclualtes the weighted integral of a quantity over the line( in my caus: the airfoil)

Now my question is thatI want to use this CFX-Post functions to directly calculate the force ( or force/lenght). Do you have any idea that this will be Lift force?


or do you have any solution how to get the CL from Cp distribuion over the airfoil?

Regards,

Thanks in advance

[ghorrocks]

You cannot integrate pressure over a line. As a line has zero area the force is zero. Extrude your polyline to make a surface, or take a cross section of your airfoil to do this.

[mohammad]

Originally Posted by ghorrocks

You cannot integrate pressure over a line. As a line has zero area the force is zero. Extrude your polyline to make a surface, or take a cross section of your airfoil to do this.

Dear Glenn,
I is possible.
for example, if you create a polyline (e.g. ployline1) then you can use the following command line in an expression:

lengthInt(pressure)@ polyline1

it gives the length integral of the pressure. Thus its dimention is ([N m^-2][ m]=[J m^-2]

However I am not sure about correctness of this integral over a close loop(airfoil).

Regards,

[ghorrocks]

I guess you are right, it returns units of N/m which is force per unit length. So getting CL from this is straight forward.

[nidhids]

Originally Posted by ghorrocks

I guess you are right, it returns units of N/m which is force per unit length. So getting CL from this is straight forward.

Hello, I'm sorry for the basic question but how do you get lift coefficient from this? I got a value that was pretty large considering my total normal and axial force values. I am trying to calculate the lift and drag using the Cp at a specific cross section of the wing. I was able to create a polyline at the spanwise location using a plane and boundary intersection. I am not sure how to calculate the lift and drag at that section in CFD post. Please help.

[ghorrocks]

Just assume the length in the third dimension is 1m, so then 1 N/m for 1m results in 1N. Then use 1N in your CD calculation.

[nidhids]

Originally Posted by ghorrocks

Just assume the length in the third dimension is 1m, so then 1 N/m for 1m results in 1N. Then use 1N in your CD calculation.

Thanks for your reply on both of my questions (here and in another post). I am having another issue. I have Pressure, and X, Y, Z - Wall Shear. To get the lift over a polyline, I would need to integrate the Pressure and Wall shear in the Z direction and then add them. Same goes for the drag in the X direction. The issue is that I don't have X Y Z values of Pressure.
If I use the same Normals X and Normals Z values (as I use for the wing), the lengthInt function gives me a results that says "undefined". This is probably because the Normals are for 3D. How do I obtain the normals for the polyline? Should I use the 'Boundary Normal on Polyline 1 X' value to calculate the Pressure along X and so on?
I looked in the manual for what this term 'Boundary Normal on Polyline 1' means but I could not find an explanation. Please would you help me?

[ghorrocks]

I do not understand your question. If you have modelled the airfoil, then just use line integrals (for lines) or force() for areas to get the forces acting on the foil. Why do you not have XYZ values of pressure?