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What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil?

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Liras

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Hi,
I'm trying to choose airfoils and establish wing dimensions (and high lift devices if needed) for my design. I have a low stall speed requirement (Vs0 = 42km/h [with flaps]), so I'm looking into high-lift airfoils with good stall and thick enough for a light spar. I'm trying to balance CLmax vs wing area. It's a single seater UL design with no more than 180kg gross weight.

I have a problem though. It's hard for me to estimate the actual, practical CLmax. My Reynolds Number is 1 milion. AirfoilTools.com says USA35A would give me some CLmax = 1.8, airfoildb.com, on the other hand, goes over 2. My wings will probably be of rectangular planform and AR~7, so I conservatively multiply the CLmax by 0.9. So, I've got: 1.62 and 1.8. That's quite a difference. My choice of high lift devices and wing area depends quite heavily on the airfoil's CLmax and 0.2 is a lot.

So, a question to the more experienced than me, which prediction is closer to the truth? Am I right to multiply by .9?

P.S.: At cruise Cl of 0.25 USA35A generates quite a bit of drag (Cd = 0.012-13).. I won't be going fast, some 120km/h but still.. maybe you have some better suggestions for airfoils for this kind of plane? Maybe a combo of 2? Just remember the stall speed req...

Thanks,
Liras
 

lr27

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

Here's a paper with info from old wind tunnel tests:
https://ntrs.nasa.gov/search.jsp?R=19930081155
Not sure how good the tunnel was. If we believe it, then 1.55, or 1.45 with a "screen" which provides some turbulence. There are some modern airfoils which are supposed to give similar performance without as much pitching moment. For instance, there's one called Ara-D 20 percent that Xfoil thinks provides similar lift with about half the pitching moment. The NACA 43018 gives up a little lift but has a very low pitching moment. There are several airfoils designated FZX.... that look promising in Xfoil.

Flaps can help quite a bit but of course that's more work. And more pitching moment.

Keep in mind that, at high angles of attack, you airspeed indicator may read low.

Do you really need 42 km/h stall? That's lower than the American part 103 ultralight regulations require!

If your design resembles the usual "ultralight" as seen here in the USA, then you may not even notice a few points of extra drag. Maybe about as much extra drag as your head, if it's exposed to the airflow. However, depending on the construction of your wing, the drag numbers may be a bit worse than you'd expect. Ever see a photograph of a fabric covered airplane in flight where the covering balloons upward on the top of the wing?

Bear in mind that I probably do NOT have more experience than you.
 

BBerson

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

VG's might help.
The Liebeck airfoil might be worth a look. Camber without flaps is probably lowest drag.
 

lr27

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

VG's might help.
The Liebeck airfoil might be worth a look. Camber without flaps is probably lowest drag.
Which Liebeck airfoil?
 

autoreply

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

Double slotted flaps?

CL of 1.55 is roughly what a modern laminar profile gets with short plain flaps. Slotted can be 80% higher.
 

Swampyankee

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

VG's might help.
The Liebeck airfoil might be worth a look. Camber without flaps is probably lowest drag.
I would be careful with Liebeck airfoils; some have terrible off-design qualities
 

BBerson

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

Which Liebeck airfoil?
I don't know. Depends on the RN and aircraft as always. Probably more suited to human powered aircraft. Somewhat controversial, I suppose.
I use my own airfoils.
 

WonderousMountain

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

I agree with both the 1.5 CLmax, and double slotted flap suggestion.

Generally the flap deploys in one motion with both foils fixed relative to each other. Single slot is not so bad, and is considered lighter usually. I mean the Delta change from flapless is good.

Also, consider under-camber, since it's an UL with a max speed restriction per regulations. There are a good variety of single element foils with respectable CLmax that get overlooked for slow flight.

Pitching moment is higher with undercamber, but it's high with flaps deployed also. Don't get over-excited about a good numbers. They're usually either false, or come with a catch.

Some foils wings are better than others though. Could you tell us weather this is hard surfaced or fabric???

LuPi
 

BBerson

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

I don't know if there is a modern airfoil for ultralights?
Looking through my Eppler airfoil book I find airfoils for the low RN human powered aircraft and sailplanes. But nothing in between. The sailplane airfoils assume cantilever and a smooth shape and are quite thick.

The ultralight airfoil would also be different if braced or cantilever. So as always, it depends.
 

lr27

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

I think Martin Hepperle has some airfoils for ultralights, but you'd want to check on that. I'm sure a search engine will find his site quickly.

You might try looking around on the UIUC airfoil coordinates database. Another place found easily with a search engine.

A typical ultralight has a relatively low aspect ratio and wide chord, so pitching moment may be more important than on other aircraft.

The less accurate your construction method is, the more skeptical you should be about max lift numbers. Also the more likely a "turbulent" airfoil may be a better choice.
 

BBerson

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Re: What would be a reasonable CLmax estimate for the USA35A thick, high-lift airfoil

I am looking at the Eppler 580 designed for the Grob G109 motorglider. Designed for high Cl Max in rough conditions (rain) from .7 million to 3 million RN. Max Cl of 1.7 at 1 million. It apparently gets high CL from the trailing edge cusp. Sort of like flaps set fixed at 10° as I see it.
I don't like the sharp leading edge and wondering about increasing the radius? The STOL kit leading edge cuffs I installed on Cessna's increased the radius dramatically for improved stall.

The MH 114 looks much like E580 with more aft camber. Or, sort of like fixed flaps at 15° degrees, as I see it. No magic really getting high lift with flaps or aft cusp or whatever it's called.

I suppose having fixed flaps with no gap is more efficient than movable flaps.
 
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