# Maximum lift values

Discussion in 'Aircraft Design / Aerodynamics / New Technology' started by orion, Jun 24, 2008.

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1. Jun 24, 2008

### orion

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This is an interesting discussion excerpt that i got this morning from one of my associates. It was written by Mark Drela:

Even so, there are some trends related to experimental data that I think we should look at. Comparing the Clmax results between Xfoil and NACA data ("Summary of Airfoil Data" by Abbott and Doenhoff) it seems to me that the dispersion between them increases for airfoils with small LE radius. Some examples:

Naca 0006 - Xfoil Clmax=1.57 / Tunnel Clmax=0.90
Naca 1408 - Xfoil Clmax=1.83 / Tunnel Clmax=1.35
Naca 0010-34 - Xfoil Clmax=1.40 / Tunnel Clmax=0.80

Naca 0012 - Xfoil Clmax=1.86 / Tunnel Clmax=1.60
Naca 2412 - Xfoil Clmax=1.94 / Tunnel Clmax=1.70
Naca 23012 - Xfoil Clmax=1.89 / Tunnel Clmax=1.80

Before blaming the wind tunnel data, maybe it would be interesting to think about whether there are any kind of simplifications or assumptions in the Xfoil code that under these conditions might lead to incorrect values.

Very few people seem to realize that _some_ of the NACA airfoil data is very strongly affected by compressibility. For the NACA 0006 at Re = 9M, these are the CLmax values I get versus Mach number:

Mach CLmax
0.00 1.56
0.05 1.53
0.10 1.46
0.15 1.33
0.20 1.21
0.25 1.06
0.30 0.94

The reason it's so sensitive is that with the very small LE radius the flow at the LE can easily become supersonic at large AoA's, even if the freestream Mach is low and the bulk of the flow is effectively incompressible. But the supersonic or nearly-supersonic flow at the LE strongly degrades the separation resistance of the downstream boundary layer, and thus decreases CLmax.

So the big question is what's the Mach number of the test data. There's no way to know, because it was never documented in the original NACA reports. The main report on the pressure tunnel which was used for these tests is here:
http://naca.central.cranfield.ac.uk/reports/1947/naca-tn-1283.pdf The models had a 2 ft chord, so the 9M data was at 4.5M/foot. The chart in Figure 10 then shows that the Mach number could have been anywhere from 0.10 to 0.30, depending on the pressurization of the tunnel. So an Xfoil calculated CLmax number could be anywhere from 1.46 to 0.94 -- pick one.

The bottom line is that one should not take the Abbott & Doenhoff CLmax numbers verbatim without considering the possible Mach number effects, especially if they have low camber and/or pointy leading edges. For airfoils with larger LE radii, like the NACA 0012, the CLmax is much less sensitive freestream Mach, but the effect is still noticeable.

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2. Jun 25, 2008

### Othman

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Thanks for sharing that little bit of information. Something to keep in mind.

3. Jun 25, 2008

### orion

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The one thing to keep in mind here though is that when considering high clmax values we must take a realistic look at the speeds that are anticipated. True, for maneuvering, it is probably realistic to use the M.3 value but generally when considering clmax, we're looking at Vmin. Even with flaps, the deltacl with flaps is added on top of the clmax of the section at the anticipated speed. As such, for GA, the most applicable value would be down at the M.1 range or lower.

4. Jun 25, 2008

### Rom

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When choosing an airfoil for my build, the CLmax wasn't the leading reason for choosing that profile , since flaps will take care of the slow flight during approach and landing. The CL as related to L/d during climb-out, the L/d during cruise and the sudden increase in drag at a given AoA with a laminar flow section (can't remember what that was called) were mitigating factors. Also How fast the CL dropped off when reaching a high AoA was also important. The GA35315 has many of those qualities. It will be interesting see how it is really going to perform in actual flight, attached to an airplane; just part of the fun.

5. Jun 25, 2008

### orion

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It should do quite well.

One of the problems of dealing with the old NACA database is that it showed certain behaviors in the wind tunnel that were not really realistic for the real world. One of these characteristics was the "laminar drag bucket" that some like to use for predicting their airplane's performance. The problem is that for real world application, even when using smooth composites, that drag bucket does not really materialize and so the rule of thumb is to simply draw a curve between the two sides of the bucket for the more realistic values.

Another issue is the "standard" roughness plot. Given the small size of the wind tunnel models, the 80 grit sandpaper they used would be equivalent to what you might get on your wing after flying through a swarm of locusts. As such, that too is generally not used.

6. Jun 25, 2008

### Rom

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It's nice to know that the airplane won't suddenly hit the brakes when hitting 8-10 degrees AoA.

7. Jun 25, 2008

### rtfm

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Hi,
Drag buckets are good things. One of the reasons I favour the 747A315 is because of its prominent drag bucket. If this (in the real world) isn't a reliable characteristic of the actual wing, then are there any compelling reasons to use this airfoil? I had not heard of the GA35315 airfoil. How do the two airfoils compare? I was working off a CLmax of 1.2 for the "clean" 747, and CLmax=2.1 in the flapped condition.

I'm a long way from actually needing to lock myself into an airfoil, but I'm curious. For my project which requires a low (42kt) stall speed and a relatively high (130kt) cruise, are there any airfoils members of this forum might recommend? I plan to use rather large split flaps along about 60% of the wing to achieve the target stall speed.

Regards,
Duncan

8. Jun 25, 2008

### orion

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I used the 747a315 on the racing bipe I did years back and the performance was great (two Gold Cup wins at Reno). We however never did the reverse testing to see how close the design numbers were (customer wasn't interested). One of the benefits of this section is that it is not so sensitive to the laminar separation bubble that is more commonly associated with the other classic laminar sections.

The GA series of sections are those developed by Harry Riblett. They get their basic shape from the classic laminar shapes but are improved at the leading edge for better flow and stall properties. Currently I use these more than anything else.

9. Jun 25, 2008

### rtfm

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Thanks for this. So it would be a good idea for me to possibly be thinking about the Riblett airfoils instead? I have read, however, that in the grand scheme of things, the actual difference between airfoil performance is relatively small, especially when one compares two essentially "good" airfoils.

Regards,
Duncan

10. Jun 25, 2008

### orion

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Very true. Often times it's the secondary things that become the selection drivers like stall quality, internal volume, etc.

11. Jul 6, 2008

### GESchwarz

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What is the big advantage of a split flap?

I can't imagine that you get a lot of extra lift with such a device. When I think of split flap, I think of the Douglas Dauntless, which is a dive bomber which uses it's split flap to generate lots of drag in a steep dive.

I would imagine a split flap would be useful in a power on approach if you want to be able to slow down quickly when you chop the throttle at touchdown.

12. Jul 6, 2008

### djschwartz

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Actually, the performance of a split flat isn't much different from a simple plain flap. Good discussions of various flap types can be found in "Theory of Wing Sections" by Abbott and Von Doenhof (sp?) and "Fluid Dynamic Lift" by Hoerner.

Dive bombers typically used a perforated flap or spoiler as a dive brake that could be deployed at high speed for increased drag to allow steep dive angles. They also have more typical flaps for increased lift for landing

13. Jul 6, 2008

### rtfm

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Hi,
Interesting - however, my understanding is that split flaps are actually quite a lot more effective than a plain flap. As the attached summary shows.

Duncan

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14. Aug 30, 2008

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