Airfoil Selection for 1/4 Scale Sailplane

[Phat2Wheel]

I posted this same question on rcuniverse.com. Thanks for all replies.

http://www.rcuniverse.com/forum/Airfoil_Selection_for_1%2F4_Scale_Sailplane/m_1284049/tm.htm

I am in the process of designing a wing for a 4M scale Schweizer 1-35. I need some advice in the airfoil selection for this sailplane. The AR is 20 and the wing has a taper ratio of .43. I've calculated the Re at root is approx 300K+, midwing is 150K-200K, and the tip is around 100K. For the approximate weight of this sailplane model, the CL is 0.4 at normal crusing speed. I've chosen the airfoils for the root and midwing section but still perplexed of how to select the tip airfoil. The family of airfoils is the HQ in which the root is HQ3512 and the mid wing is HQ3012. However, the HQ drag polars do not look very good at low Re's.

Through literature search, I found that some German scale sailplanes use high Clmax airfoils the tips. Tip airfoils have greater t/c than even at the root. I think this aids handling, turning for such large models especially closer to the ground. Scale flying characteristics seem to be the drivers rather than pure efficiency. I also looked at the Algea 3M wing. The tip airfoil has lower lower drag than the root at low Re but it also has lower Clmax. This is opposite to the scale saiplane in which this design seeks for the ultimate thermalling efficiency.

The rationale for low drag polar is high AR wing has high skin friction drag thus the airfoils should be optimized along the span. However, high AR wing can be proned to tip stalling in high bank turns. Therefore, high Clmax airfoil should be used in the tip section with the expense in additional drag.

The question is should I choose a tip airfoil for the low Re drag polar characteristic or should I choose to go with high Clmax? How should I resolve this compromise? Thanks.

[Ollie]

The usual requirements for a good wing design are:

1. High coefficient of lift at slow speeds to reduce landing speed with adequate stall margin.

2. Low induced drag through high aspect ratio and an elliptical lift distribution.

3. Tip stall margin by adequate maximum lift coefficient and washout.

4. Low drag at high speeds and low lift coefficients for good wind penetration.

5. Enough airfoil thickness for a strong and stiff spar.

6. Gentle stall characteristics.

I highly recommend the S4233 because of its high maximum coefficient of lift, its low drag over a very wide range of coefficients of lift, its thickness for a strong stiff spar and its very forgiving stall characteristics. There are better air foils for particular aspects of the objectives but the S4233 is one of the best compromises among all the objectives.

[SZD16]

I agree with Ollie 100% The 4233 is an excellent scale airfoil......very good preformance....but the best part is it is 13.66% thick which works real well for vintage scale.......

[Phat2Wheel]

Ollie

Thanks for the suggestion of the S4233 airfoil. My first instinct is to select a thick airfoil with a good amount of camber. From what I've read scale sailplanes tend to have higher wingloading than the norm. I've just computed the root bending moment and found that it is aroung 1500lb-in. Wow a big number so any additional thickness will help. I also want camber since this helps with the lift coefficient. I made some run with XFoil to compare with my current airfoil selection: HQ3512 (3.5%camber, 12% t/c). I also added the HQ3513 since this one also has 13% t/c similar to the S4233.

The Clmax for the HQ series airfoils are from 1.35 to 1.45 (HQ3513 has the higher number). The s4233 Clmax is at 1.25.
Cl at 0alfa are 0.47 for both HQs while S4233 has 0.34. So at least for the root airfoil, I will choose the HQ series. I will post the sceen dump of Xfoil once I find out how to do it.

That being said, the question remains as what tip airfoil should I use. I reread John Stackhouse explanation on aerodynamics of turns. He said that the inside wingtip will have higher alfa especially in tight turn. I assume that low Clmax is not suitable since the wing will tipstall. Ollie mentioned that washout will also help. This option may decrease cruise performane.

[schrederman]

I'd sure like plans for this model. I have 125 hours in the real thing.

Jack Womack

[Ollie]

The HQ 3.5 sections look very good down to a Cl of 0.2 and the higher maximum Cl is an advantage.

The problem is that the high aspect ratio and the small taper ratio make it harder to get both a reasonable tip stall margin and an efficient lift distribution over a wide range of angles of attack. See:
http://aero.stanford.edu/WingCalc.html
One possible solution is to decrease the camber toward the wing root and retain the 3.5% camber at the tip. This will introduce aerodynamic wash in which can be compensated by lots of geometric wash out. Use a camber of about 1.5 % and a thickness of 13% at the root. At the tip, use a camber of 3.5 % and a thickness of 8%. The geometric wash out should equal the difference in zero lift angles of attack between the root and tip airfoils plus a degree or two.

The loss of over all maximum lift coefficient of the wing could be compensated by a lighter wing loading if you are willing to carefully select your materials and refine the structural design.

It sure isn't easy to meet all performance and handling objectives when the planform is constrained by scale considerations.

[Phat2Wheel]

I haven't given up on the S4233 airfoil yet. Since this airfoil has "better" looking drag polar, I tried to modified so I could use it at the tip. So the final number are: 12%t/c, 3.5%camber. I moved the max t/c from 33% to 29%. The Clmax is now up about 1.4 and the drag polar looks alot "better". One thing I don't understand is Clmax is higher at lower Re. This is counter intuitive but I think the high Clmax is suspected due to the low Re. So this is what I will use in the wing: HQ3512root - HQ3012 mid wing - S4233mod at tip. Ofcourse these will change once I find out how much the fuselage weights. I just finished making the frames for the fuselage mold master. Construction begins next week.

[Phat2Wheel]

I was having trouble with this pics somehow.

[Ollie]

Referring to Soartech #8 and the wind tunnel tests of the S4233:

The wind tunnel test model had an average contour error of 0.0043 on a 12 inch chord. The peak errors were at the leading and trailing edges and, were about 0.03 inches.

The wind tunnel test models had measured maximum lift coefficients of 1.28 at a reynolds number of 101,000.

The S4233 was also tested with a turbulator strip at 20% chord. The maximum lift coefficient was the same as without the turbulation strip. The measurements of the turbulated airfoil match your XFoil results very closely except for the lower maximum lift coefficient.

The discrepancy in maximum lift coefficient might be explained away by the contour error at the leading edge of the wind tunnel test model. If that were the cause of the discrepancy, then it raises the question of how accurately the airfoil must be built to achieve the theoretical performance expected. If the tip chord of the model were less than six inches then an error at the leading edge of less than 1/64 of an inch might be enough to lower the maximum coefficient of lift some 8%.

When selecting the washout for tip stall margin it might be worthwhile to consider an allowance for construction contour errors.

[Purdue Aero Man]

Originally posted by Phat2Wheel
One thing I don't understand is Clmax is higher at lower Re. This is counter intuitive but I think the high Clmax is suspected due to the low Re

Phat2WHeel,
Perhaps the explanation for the increase in Cl_max as the Re is lowered is due to less flow separation about the airfoil. At low Re, depending on the shape of the body, flow can actually remaim attached longer, which in this case may be increasing your Cl as there is less turbulent flow and separation around your airfoil. I'd also like to caution you about relying on XFoil at too low an Re. From your attached polars, the Re's you are working at are in the 100,000 to 300,000 range. XFoil will not generate accurate data at such low Re. Also do not rely on XFoil for your pitching moment calculations as it cannot accurately predict the correct lift distribution along the chord.

If you are in fact looking at operating at such a low Re, a sight that has a lot of different polars and coordinates for airfoils is http://www.nasg.com/index-e.html

[Phat2Wheel]

Purdue Aero Man,

Thanks for the website. I checked on the S4233 airfoil and Clmax is the same for a Re ranging from 100K to 300K. So at least Xfoil is not predicting this airfoil correctly at Re300K. Cl plot for 100K look similar. I also checked Helmut website for the HQ airfoil data. Seems like my Xfoil runs give similar results to his posted data. However, I suspect he was using some types of analysis tool instead of real test data. I have to translate from German to know for sure.

I will continue my contruction discussion on the "Scale Sailplane" forum since it is more appropriate there.

Phat