Aerofoil pressure drag calculation
Hi everyone, this being my first post I will take this opportunity to say that you seem to have a very interesting forum going here!
I have a question with regards the calculation of pressure drag over an aerofoil (or airfoil if you prefer).
I am using a panel method coupled with an inverse boundary layer solver to calculate the velocity distribution and boundary layer profiles. The inverse boundary layer solver accounts for interaction of the boundary layer with the inviscid outer flow and can calculate separation bubbles. It gives as an output the modified pressure distribution as a result of boundary layer growth (and separation where it occurs).
In order to calculate the pressure drag coefficient I resolve the pressure coefficient term at each station into its axial component and integrate along the chord. For each station (i) this is done using:-
Cd_p(i) = Cp(i) * dy(i)
I then sum these terms over the upper surface, and do the same for the lower surface; the total pressure drag coefficient is then giving by adding the upper surface component to the lower surface component.
The problem that I am having is that the upper surface contribution is giving a negative pressure drag term, and is of larger magnitude than the lower surface term (which is positive) giving me the clearly nonsensical result of negative pressure drag.
This problem is due to the terms near the leading edge. This seemed counter-intuitive to me as pressure drag arises due to boundary layer growth and separation which are effects that only become significant in the region of the trailing edge.
Any suggestions are much appreciated.
Thank you for reading and apologies for the long post...