Has anyone out there ever analyzed apllication such as high lift devices on the wing (to be more specific I want to carry out 2D multiple element airfoil types of analysis to estimate CL max & CD max)?

For the past few days, I was trying to model a NACA 2415 airfoil (clean airfoil w/o any slat or flap, RE = 6e6) at alpha = 0 deg and 2 deg and compare the Cl and Cd values with those in the Theory of Wing Sections. I have carried out the mesh dependency test on the same test case and at the end, the Cl predicted is only about 50% of the theoretical value, and Cd is 300% of those in the book. My mesh density on the airfoil surface is (0.5%*airfoil chord). Is this enough?

Can anyone give me some insight on how to tackle this type of problem? Greatly appreciated the help Thanks!!!!

If you're doing a single element airfoil, use XFoil (it's free) to at least get reference values, so that you know where you are slipping up. Averaged 0.5% of chord for cell-spacing is plenty fine enough, but you probably want to cluster it more to the leading edge. Drag will always be difficult to get right, because the RANS turbulence modeling approach will not automatically predict laminar - turbulent transition correctly. If you are using Fluent or CFD-Fastran I can give you example files that work OK.

Thanks for your insight. Currently I am running CFdesign (a finite element code).

By the way, when you say you have some "ok" example files, what do you mean by that? Have you ever obtained any sorts of results that are within 10% of theoretical values?

Getting lift within 10% is no problem. The drag can be whatever you want it to be ;-) , because it is dependent on selection of turbulence model, enforced (or not) transition, wall functions, normal spacing at the wall, etc .... If you really want to study 2D multi-element wing profiles, perhaps MSES (big brother of XFoil) is the code to use. That will give you far better drag values.

I have been looking for MSES all over the internet but I couldn't find any .exe files. Is this MSES free like Xfoil (I do have Xfoil).

I know it is very difficult to get the correct Cd. But I also have problem getting the lift. Right now I am trying to model NASA LS0417 airfoil (in NASA TN D 7428). I am starting the modeling of the airfoil from 0 degree up to 16 degree alpha. At high alpha case, I can only get about 50% of the lift because the high pressure gradient at leading edge suction side is so grossly underpredicted. All my pressure side Cp vs x/c on the other hand has good agreement with the paper. Anyone out there who has experience in this type of modelling please give me some help. Besides keep on increasing the mesh count right around the leading edge there seems to be nothing else I can do, but the suction side pressure is still under estimated by alot.

Thanks for the help in advance.

MSES is not free, except possibly for academic research purposes. Best plan is to contact Mark Drela (author of the program) at MIT directly. Are you at least getting the lift at zero angle of attack and the lift curve slope at small angles of attack right? At high angles of attack the choice of turbulence model may become quite important, the standard k-epsilon model would typically be a bad choice!

At this point, for alpha = 0 degree, I can only get about 50% of the Cl (pretty much the same as alpha = 16 deg case, about 50-60% Cl). So what are the catches for this type of problem. Before modelling this NASA LS airfoil, i also tried to model NACA 2415 and clark Y airfoil, and I get almost perfect match at low alpha and about 90% Cl at alpha near stall. But for this new NASA airfoil (not really new, 1973) I am really puzzled at the result. Can you please give me additional guildlines?

by the way, I am aware of the fact that standard k-epsilon can't deal with recirculation well, that's why it is a bad choice for high alpha cases, but at low alpha cases where there is no separation at the suction side or trailing edge, do you think k-e still good?

You may want to focus some attention on the trailing edge area rather. The LS(1)417 is a slightly bizarre airfoil, in the sense that it is very heavily aft loaded, and if you don't get the flow at the TE right you won't get the "suction" right either. That thick trailing edge needs to be modelled as well. This is to me an interesting topic, but maybe it should be continued by e-mail. Contact me at charles.crosby "at" kentron.co.za , replace the "at" symbol with the usual @ (trying to avoid spam here ;-) )

[cfd.aero.soam]

hi Charles crosby

I am Prashant from India. I am working in National Aerospace Laboratory. I have created mesh over multielement airfoil in Gambit and now i am using Open Foam as solver. At M=.2 and 0 AOA. In solving I dont knoew hou to converge solution .Please give me your guidence and some tutorials ,which i can follow in this regard.
Prashant Kumar
cfd.aero.soam@gmail.com

[cfd.aero.soam]

hi Rookie

I am Prashant from India. I am working in National Aerospace Laboratory. I have created mesh over multielement airfoil in Gambit and now i am using Open Foam as solver. At M=.2 and 0 AOA. In solving I dont knoew hou to converge solution .Please give me your guidence and some tutorials ,which i can follow in this regard.
Prashant Kumar
cfd.aero.soam@gmail.com

[wanfuhh]

hello sir, can i have the example of file for analysis using fluent that u said before? i would really appreciate it. here is my email. wansyahmi17@gmail.com

[Totalsim]

This may sounds silly but have you represented the geometry accurately? It may be something as simple as the blockage ratio.

[NaderNekoubin]

you may have three possible mistakes. 1: solving the main equation including of Navier-Stokes and energy equations wrongly. in this condition, you can check the whole procedure. 2: calculating the cl and cv wrongly. in this condition, you can first check the contours for velocity and pressure around the airfoil with experimental and numerical results in the literature to ensure about the solution; then you can check the formulas for cl and cd calculation. 3: the quality of used grid. it has little possibility, but using a very "bad" grid can cause such deviation from the actual and valid results. check the mesh!