NACA 6-series cusp

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wsimpso1

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Riblett simply adds the thickness and camber together. NACA takes the camber line and then add the thickness to that curve. Riblett's book has a good explanation of the two methods.
When I am back for a couple days and caught up on more important stuff, I will run NACA and Riblett profiles, take apart the mean lines and thickness distributions and combination methods, and see just how big (how small?) the differences are.
 

pictsidhe

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I don't think the difference is very big at the moderate camber that I am using. At some point, I'll add the NACA method to my spreadsheet and find out.
 

Starflight

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Thank you so much "pictsidhe". Nice to see it again. I also looked at the NACA 63-614 a=0.4 when I had access to Javafoil and found their Cl/Cd curves were gently rounded turbulent style compared to the sharper bends of the pure laminars. My minimum standard for Clmax at Reynold's numbers between 1.4E to 2.6E, has to be more than Cl = 1.5 Both of these airfoils met that spec. with ease. Now, if we could just get them into a wind tunnel :)
 

PiperCruisin

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I have not noticed an issue with the cusped airfoils such in the plot in post #20. Javafoil seems to do a fine job. The last 3 points were as listed.
upload_2019-5-8_8-6-17.png
 

PiperCruisin

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Thanks for airfoil coordinate plot.
If you are interested in other NACA 6A series airfoil coordinates here is a link http://pdas.com/sections6a.html
The PDAS.com website has a lot of cool programs.
I found that site once and then lost it. Thanks. I plugged in a 64-A415. The TE looks better, but the cd still seems high in Javafoil, but not in XFLR5...hmmm.

Maybe has to do with roughness settings?
 

PiperCruisin

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Neither xflr5 nor javafoil are very accurate. See this page
What would you recommend as an alternative? XFoil direct or pay for something?

I have noticed some inconsistency that Javafoil seems to give some good results, but drag can be predicted much higher and I thought the xflr5/xfoil results were too choppy to interpret.
 

wsimpso1

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I plugged in the numbers for the NACA642A215 and Riblett 37A215 foils, and plotted them in the attached Excel file. The big differences are that Harry's foil has a tad more camber and with that a somewhat steeper midline apparent at the leading edge. The former is a bit of a surprise, the latter is Harry's big point...

I even tried to use NACA's combination method for assembling the NACA foil. Lines up pretty darn close to the NACA coordinates, so I guess I did it right.

Looks like the combination method when used in combination with Harry's adjusted mean line makes tiny differences to the finished shape. We could check the mean lines and see how much difference that makes...

Billski
 

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pictsidhe

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What would you recommend as an alternative? XFoil direct or pay for something?

I have noticed some inconsistency that Javafoil seems to give some good results, but drag can be predicted much higher and I thought the xflr5/xfoil results were too choppy to interpret.
xflr5 uses the xfoil code. I don't know of anything significantly better without a significant price tag...

I plugged in the numbers for the NACA642A215 and Riblett 37A215 foils, and plotted them in the attached Excel file. The big differences are that Harry's foil has a tad more comber and with that a somewhat steeper midline apparent at the leading edge. The former is a bit of a surprise, the latter is Harry's big point... I even tried to use NACA's combination method for assembling the NACA foil. Lines up pretty darn close to the NACA coordinates, so I guess I did it right.
Right, I'm stealing that!
 

Retiree

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Neither xflr5 nor javafoil are very accurate. See this page
This points up the weakness and strength of these empirical methods. They only do a fair job of predicting actual numbers. Their strength is comparing similar airfoils. So if you have a good set of data on an airfoil similar to the one you want some numbers for, you can estimate the numbers of the comparison airfoil.
The worst estimate is usually the maximum lift coefficient. The reason is the codes do not model the physics of separated flow. A number of people have come up with a rule of thumb to get a better estimate of the max lift coeff. When the flow has separated from 20% off the end of the foil or the drag has reached .0250 the airfoil has reached its maximum lift coefficient. I do not have a reference for this.
 

PiperCruisin

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Right now it seems the data varies a lot depending on what I am using. For the sake of sanity, maybe I'll just pick one and pretend it is accurate. One I start to like is a NACA 63-417 a.5

I suppose I could transition to a NACA 63-A415 to reduce the cusp in the aileron section if it is a problem.

In Javafoil there can be a large difference when using "smooth" vs. "NACA Standard" for roughness depending on the airfoil. The 93 series seems to give more reasonable numbers with standard roughness while being optimistic with smooth...or not. Who knows? I'm mechanical so anything over 2 to 3 significant digits and my eyes glaze over.
 

DeepStall

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In Javafoil there can be a large difference when using "smooth" vs. "NACA Standard" for roughness depending on the airfoil. The 93 series seems to give more reasonable numbers with standard roughness while being optimistic with smooth...or not. Who knows? I'm mechanical so anything over 2 to 3 significant digits and my eyes glaze over.
NACA Standard Roughness for their 24" chord models was 0.011" tall grit over the leading 8%, 5-10% coverage per area. That grit height/density is very severe when scaled up to a vehicle in normal service. (well ok maybe if you never clean the bugs off) Again suggest finding another foil with good wind tunnel data available to calibrate your computational techniques against, then applying that technique to the foils under consideration.
 
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