<table cellspacing="0" cellpadding="6" border="0" align="center" id="post5022434" class="tborder" style="width: 1028px; height: 121px;"> <tbody> <tr valign="top"> <td class="alt1" id="td_post_5022434" style="border-right: 1px solid rgb(0, 0, 99);"> <div id="post_message_5022434">Hello,

I've been repeating some basics and I'm stuck. As far as I remember, the maximum AOA allowance for a given wing, has been defined with its chord length. That goes for either highly swept wings like Mirage delta, or straight ones like the Spitfire's, so it isn't directly related to high-sweep leading edge vortex or aspect ratio.
IIRC, it had something to do with movement of center of pressure (cp) along the chord line. So, again IIRC, the stall occurs (in 2D sense) when the cp departs the wing's area. Longer chorded wings have this travel space larger than short chorded, hence larger AoA limit for them.
Can anyone verify that? and if I'm all wrong, please post a link or answer with correct explanation.
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I hope the information on this website will help you:

http://www.mh-aerotools.de/airfoils/index.htm

Regards!

The center of pressure travel concept was dropped years ago in favour of a fixed airfoil "pivot point" at the 25% MAC chord point and a "pitching moment" amount to describe the pitching effects that were previously covered by moving a virtual balance point which was the center of pressure concept.

This was done, near as I can make out, because the Center of Pressure concept went into the toilet for the specific lift value of CL=0 where the CP was located infinitly far behind the trailing edge. Mathematician's may not mind infinities in their equations but engineers hate them. Hence they came up with the idea of a consistent pivot point and the pitching moment which avoided the whole irritating infinity issue.

BUt all else being equal you're right. The stall angle is somewhat affected by the chord. But I do not think you can lump in acute angle delta and strongly swept planforms with constant chord wings. Strong sweep angles do generate strong leading edge vortex formations that muddy the whole stall issue to the point that you can't really compare them fairly.

Guys, thx a lot for helping.

Phew...I've been through a good two dozens of sites and I think, I got somewhere. So, it turns out that fixed CP concept works for symmetrical, but not cambered wings. Further, it also seems, that in reality the CP for CL=0 isn't in the infinity, but just outside the trailing edge, meaning for zero lift as the CP departed lifting area and that's the correlation with chord I was looking for.
Well, I remembered it was working something like this, but wasn't sure exactly how. It's good though to use movable CP in sense of proving AOA limit depends on chord length, because it's easy to remember, being visually logical concept.
One other thing. For delta wings, the stall occures more gentle than with trapeziodal ones. With the "traveling" CP, it's easily explainable, because as the CP moves forward, it reduces lifting area from the tip towards the root, by exiting the wing's lift area (going ahead of leading edge). OTOH, in trapezoidal wing, the CP is either inside the wing's lifting area, or is out, so the transition is more violent and that corresponds with high AR wings having steeper Cl/AOA curve than lower AR wings.
Ok, I hope this makes sense.

Cheers,Cola

Hope nobody minds if I don't buy any of the answers .
If a wing has a round shape like a barrel then it won't stall at any angle irrespective of what the chord (dia) may be
(Also it won't lift worth a darn)
If we take a big hammer and start pounding the barrel to change thickness to chord ratio
the angle at which it stalls will progressively change
once we have pounded the barrel flat - it will be great at lift but super critical at operating angles
There you have have it
In my own book of airfoils , there are two basic shapes:
round and flat
all the rest are simply compromises which are hopefully matched to a specific requirement.
no best ones no worst ones just lots of em

Sorry, but where did you get that idea?

The CP doesn't move the lifting area with it at all.

True.
But, let's consider the delta wing with:
1 meter of root length
1 meter of trailing edge length
1.41 meter of leading edge length.
The CP is a lift resultant, so it's actually situated on longitudinal (chord) axis and moves forward and afterward, depending on AOA. Now, if we take a look at this delta wing from top projection, we'll see that the best spanwise lift distribution is when the CP is way backward (small AOA). As we increase AOA and move CP forward, the effective span is reducing (CP for each wing is moving forward, but towards the wing root, since CP can't be outside the lifting area) and so is spanwise lift distribution. However, the AOA increase produces higher overall lift increase (and drag), than the spanwise lift distribution reduces that same lift.
In the end, we get a gentler Cl/AoA curve, than that of the straight wing and that corresponds with empirical data.
Hm...does it makes more sense now?

No

And it's based on assumption that's not proven here or elsewhere. For example, your idea that increasing the AOA supposedly reducing the effective span or the same result from movement of the CP either forward or back.

The idea that changing the AOA reduces the lifting area seems to be your original idea. It doesn't happen.

That decision has always seemed somewhat humorous to me.

If you think about it, why exactly would it matter where no lift was. Zero lift produces no pitching moment after all. So why would it matter where the zero value was.

What did seem true about their being able to arbitrarily decide "where" showed how little value there was in their proofs.

[img]file:///C:/t/moz-screenshot.jpg[/img][img]file:///C:/t/moz-screenshot-1.jpg[/img][img]file:///C:/t/moz-screenshot-2.jpg[/img][img]file:///C:/t/moz-screenshot-3.jpg[/img] Thx, for bearing with me...
Now, check this shematics.
CP moves until aerodynamic center and that's the furthest point the CP will go.
After that, CP begins to move toward the stall AoA, after which departs wing's chord, resulting in zero lift again (like in the start position).
Now, observe tirquise line (CP) while it leaves wingtip's chord to reach aerodynamic center. You see, that CP's line of lift, will be cropped by leading edge on delta wing. On the straight wing, this doesn't occurs in straight wing, hence the steeper Cl/AoA curve and better spanwise lift distribution, but lower max AoA.
Finally, observe how CP must travel a longer distance in delta wing to reach AC, than in straight one. The pure arrow (Me262, Sabre) wing will be the worst of all, because it'll feature short MAC, but high sweep and therefore large tip moments and a demanding structural integrity to keep it together, while having steep lift curve and be easy to overstress.
Am I correct?
In any case, thanks a lot guys.

Yeah, it seemed like sort of a funny reason to dump the system. But engineers need to be able to describe their stuff mathematically and that infinity issue is not something that makes for clean math. By switching to a convention that puts the lift center at the 25% chord point and then just deals with the "torque" by describing it as a pitching moment at worst the value will go to 0. And that's something that works out a lot better for calculations of any sort.

Either system works. After all it's just a way of describing how the wing is producing it's lift. The convention for describing it that works best with the other aspects that need to be considered is obviously going to be the cleanest one.

Where are the guys like Ben Lanterman and Bank2Turn when you need their input? They worked in aerodynamics engineering and could better explain some of the reasons for the shift. I'm pretty much parroting what they have said in the past about this but I may be getting some of it wrong.

Cola, I think you may not have some of your Cp locations quite right. It's been a while but I seem to remember that the Center of Pressure tended to move forward with increasing angle of attack until at the stall it was at the leading edge or perhaps it was said to "fall off the front" of the leading edge at the stall point or such thing. I may be mis-remembering this aspect. I never really looked at all this stuff seriously until the fixed life locus and pitching moment was being used and that's the one I learned to understand.

Why is the pitching moment idea better? Because since the pitching moment is a torque it is easier to look at a torque from the wing and a torque (lift) from the horizontal tail and understand that the stabilizer will be able to deal with the wing's behaviour or not.

Yes, I sort of remembered that too, but I wasn't sure.
However, it seems the reflex-cambered (with upwash) wings have CP in infinity, but forward of the wing, so as the AoA increases, the CP acually moves backwards until it reaches AC and then back forward to infinity. I'd guess that in reality, for zero lift, the CP is situated just in front of the leading edge.
As for the "falling of the front", it seemed logical to me too, but when you think better, you'll see that it's impossible (I think, at least), because when the wing stalls, it actually doesn't get zero lift immediately (and how would you get lift from a wing whose CP is outside its lifting area?), but depending on the wing type, looses Cl gradually and during that time CP moves back to original zero lift position, but this time with AoA 90&deg; or so.
Why is a stall point behind AC? I think because, AC is a point where Lift/Drag coefficients meet and thus wing has the best effectivness. Stall point is beyond, because wing has some more chord to produce more lift, but beyond AC point drag rises exponentially to AoA.

But now, the more I think of what you said, the more I believe it was something like that. I mean I think I remember something like CP falling from the leading edge, but maybe those were upwashed (stable) wings or something.

As for fixed CP, I'd agree it's easier for calculations and I'd probably use it myself if I was to calculate an aircraft. However, I find this "traveling" CP explanation way better explainatory in terms of chordwise dependance. It's less abstract and proves direct relationship between chord and max.AoA.

Any finally, why does CP moves at all?
I think it's because of lift/drag resultant and the cambering. Cambered wing is unstable, with pronounced pitching moments (either up or down). So, as the AoA raises, the CP moves forward because lift vector is getting larger and larger and more parallel with chord. After AC, the drag component becomes dominant and results in CP's moving backwards.