Drag data - does this look right?

[joemadeus]

I'm coming up with my first design and trying to estimate drag, at least profile drag since it seems like that estimation ought to be easy. I have some screenshots from an excel worksheet I've put together with the numbers from the model -- the $10k question is, do they look correct? I'd like to know I'm in the ballpark :)

Some pertinent info:

Airfoil is Go417-A (described as, "a flat plate with camber"), 3% thickness, 6% camber (see the NASG entry here.) (http://www.nasg.com/afdb/show-airfoil-e.phtml?id=547)
The lift & drag data below was generated by XFoil 6.94, viscous mode, Re=57k, ncrit=9. The calcs converged for all alphas except a=1deg. I've read that XFoil doesn't do so well with low Re's -- but I don't know what else to go on. The Re=57k Cl/Cd numbers (calculated) are within ~10% +/- of the Rd=59k numbers (measured) on the NASG site.
The planform correction (k) was set to 1.1 when calc'ing Cd interference
Design speed is 25mph, or ~11 m/s
Weight is ~4oz, or 118g
I'm not going to ask it to do acrobatics. (at least, not TOO often ;))


A table is shown for 7.4cm chord, 30 and 36cm spans; I've also graphed this data (X = velocity, Y = grams), plus 5.0cm chord, 30 and 36cm spans (which has to fly a bit too fast for what I'm looking for.) The Cl in the table is what's needed to support a 118g model at 11.2m/s; from my estimation, I have plenty of leeway when it comes to how fast I have to fly to make it work.

I've also posted the polar for the airfoil for Re=57k.

Like I said, I'd like to know if I'm in the ballpark. My assumption is that I can estimate profile drag on the horizontal stab next, add some factor for parasite drag, then pick a suitable motor & prop.

Thanks for the help ;)


-j

[Ollie]

This looks about right to me. This airfoil is OK for FF with narrow range of use of Cl.

A R/C airfoil would better with lower camber for a using wider range of Cl but with lower max Cl.

[joemadeus]

Ollie,

So, if I understand you correctly, go with a reflexed camber, and reduce the camber a bit, and I should wind up with wider operating margins?


-j

[Ollie]

For airfoils see:
http://www.charlesriverrc.org/articles/drela-airfoilshop/markdrela-ag-ht-airfoils.
I suggest ag16, -17, -18 and -19. The ag's much better wide range (Cl=0.1 to Cl=0.9 & Cd=0.03 to Cd=0.007) of Cl/Cd compared to the 417 airfoil.

[joemadeus]

So, I finally got around to re-running XFoil and crunching some numbers -- it's been a busy week -- and the difference between the two airfoils is pretty startling. It looks like there's 30% less total drag generated by the wing, that drag doesn't ramp up as drastically as you increase speed, and there's still more than enough lift to support the weight of the plane (at 7.4cm x 36cm, 25mph).

I imagine that this is what years of experience get you, yes?

I included two charts comparing total drag vs. velocity (X = velocity in m/s, Y = drag in grams -- sorry, I seem to be all metric these days) and graphing the polars for the two airfoils.

Thanks for the help, Ollie!


-j

[Ollie]

Not my credit but to Dr. Mark Drela. He wrote X-foil. He designed AG airfoils. He is a master craftman model builder. He has better model designs that I haved see in 68 years. I have learned from Dr. Drela than all combinded designers.