Hey,

i am interested to know by setting an area of the aerofoils of a rear wing value in reference areas does that at all interact with calculating the pressure coefficient of say of an 3d aerofoil section.

I noticed that i haven't touch the reference values and i know that when i calculate lift an drag which i will do from the pressure coefficents i have to add the area into it however do i need to do anything to the pressure coeffient (i.e.) incorporate area.

as cp= p -p_inf / (0.5 x density x velocity^2)

also i am only modelling half the rear wing so does that mean that when i calculate lift and drag i only use the area of half the wing (i.e. the two aerofoil elements) and is the area just the reference area of the aerofoil shapes i.e. span x chord. or is it the frontal area, any help would be greatly appreciated. very much so.

cheers joe bryant

Hi Joe! the surface area you should use is the HALF planform surface of the full-wing since you're using half of the real model. The cp value doesn't depend on the reference area. Only the drag and the lift coefficient depend since they come from the sum of the product betwen the element area and its cp (projected then along the direction you specificed; for ex.(1,0,0)). CL=net AERO_FORCE/(0.5 density vel^2 S) I hope to be clear. Luca

Thank you so much for your help its much appreciated.

Cheers Joe Ps. will put you as an acknowledgement in my 4th year thesis for aeronautical engineering Sydney University

cheers once again

You're welcome...if you had any problem with Fluent just write!I'd be glad to give you some suggestions (if I can...). Luca

[shreyas]

Originally Posted by Luca Cavagna
;115554

You're welcome...if you had any problem with Fluent just write!I'd be glad to give you some suggestions (if I can...). Luca

Hey luca...how do I get the values of cp in the right range...I get it in the range of 1e6 etc...I need to compare it with theoretical results