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ACHIEVING THE BEST REFLEXED AIRFOILS FOR FLYING WING USE IN THE SMALL PLANE CATEGORIES

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Norman

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Ideally we want the CM to be positive at all times whilst maintaining the greatest lift and having the best area under the LD curve we can get.
The slope of the Cm curve is just an artifact of the program using 0.25c as a point of reference. Xfoil plots Cm relative to the 1/4 chord but that's not necessarily where the aerodynamic center is. To plot a moment you have to have a center point and thin airfoil theory says the AC is at 25%c so that's what most of the programs use but when your airfoil is more than a few % thick it's often not true. There is a simple formula to find the true aerodynamic center from the slope of the linear section. I posted it on HBA some time ago. I'll try to find it tomorrow. Anyway the point is that the only Cm value that matters is the Cm at zero lift (Cm0). As long as Cm0= some positive number and the static margin set to produce an equal negative moment the plane will be in trim.
 
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WINGITIS

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The slope of the Cm curve is just an artifact of the program using 0.25c as a point of reference. Xfoil plots Cm relative to the 1/4 chord but that's not necessarily where the aerodynamic center is. To plot a moment you have to have a center point and thin airfoil theory says the AC is at 25%c so that's what most of the programs use but when your airfoil is more than a few % thick it's often not true. There is a simple formula to find the true aerodynamic center from the slope if the linear section. I posted it on HBA some time ago. I'll try to find it tomorrow. Anyway the point is that the only Cm value that matters is the Cm at zero lift (Cm0). As long as Cm0= some positive number and the static margin set to produce an equal negative moment the plane will be in trim.
I look forward to hearing more on this.

Cheers
 

WINGITIS

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Well I'm blessed with insomnia so I looked up that old post now. The numbers can be found in the text ouput of XFLR5 or lifted from the Cm over alpha graph if you have coordinate output turned on or you can read them in the OpPoint window.
The CM displayed in the OP-POINT WINDOW WITH ANGLE OF ATTACK is THE SAME as on the normal CM/ALPHA screen for all the AOA?

I checked a few airfoils including checking Forced transitions set to 0 and 1?????????

It also matches the exported polars!!!
 

Bille Floyd

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The theory is good. But actually twisting a wing in flight is no small order.
Here's how it works :
The moment i drop the flap by any amount ; i get twist. The flap is on
the inboard 1/3 span ; when it goes down , the inboard 1/3 span, is now
at a higher AOA than the tip, so the inboard part of the wing, is gonna
stall first .
The designer , (Felix Ruehle) recommended a minimum 5-deg down
flap , in turbulent air, (for the Exxtacy glider) ; it's because ya get 5-deg
more twist,than that which was built in the frame, when you do that, which results in a more stable flying wing.
The Atos glider, got around that problem, by adding a small aria tail.

I had this exact same conversation with Irv Culver , in the late 80's ; but
my question to him, (at the time) was what happens when the % cord flap
is changed , and how does that affects the twist distribution, on a flying wing.


Bille
 
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pictsidhe

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If you use a zero Cm foil for a plank, you would have to fly with up flapevator. Making it a +ve Cm airfoil, but not optimised for that...
 

pictsidhe

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Here's how it works :
The moment i drop the flap by any amount ; i get twist. The flap is on
the inboard 1/3 span ; when it goes down , the inboard 1/3 span, is now
at a higher AOA than the tip, so the inboard part of the wing, is gonna
stall first .
The designer , (Felix Ruehle) recommended a minimum 5-deg down
flap , in turbulent air, (for the Exxtacy glider) ; it's because ya get 5-deg
more twist,than that which was built in the frame, when you do that, which results in a more stable flying wing.
The Atos glider, got around that problem, by adding a small aria tail.

I had this exact same conversation with Irv Culver , in the late 80's ; but
my question to him, (at the time) was what happens when the % cord flap
is changed , and how does that affects the twist distribution, on a flying wing.


Bille
I think of twist as a physical twist of the wing. Using inboard and outboard flaps to alter the lift distribution is a viable and tested way to trim a swept wing. Hortens used it exclusively from II onwards for both trim and yaw.
If you pick the right planform and flap shape, you can get a fairly good approximation of your desired lift distribution over a wide speed range. The penalty is that Cl varies somewhat over the wing. If induced drag is your issue, it's probably the best solution.
 

Bille Floyd

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...
The penalty is that Cl varies somewhat over the wing. If induced drag is your issue, it's probably the best solution.
That varying CL, using different cord % flap ; is what the conversation
with Irv Culver was all about . The flap on my flying wing is tapered
so the CL is different at each , incremental span-wise location.

Bille
 

Norman

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The CM displayed in the OP-POINT WINDOW WITH ANGLE OF ATTACK is THE SAME as on the normal CM/ALPHA screen for all the AOA?
Yes, you can get the same numbers in the OpPoint window that are used to draw the Cm over alpha graph or any other graph. I scribbled notes on the screen shots in post #26. You choose an AoA from the drop down menu at the top right and data for that AoA is displayed in the text block at the bottom right. You may have to move the airfoil to read the numbers you want.
 

Norman

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I made some new airfoils this morning that may be suitable for planks using the Fauvel 14% and the NM-66Qt that I posted yesterday. They all have higher CLmax than the Fauvel and lower surface forced transition to smooth out the Cm curve doesn't affect their lift like it does the Fauvel.
 

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WINGITIS

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I made some new airfoils this morning that may be suitable for planks using the Fauvel 14% and the NM-66Qt that I posted yesterday. They all have higher CLmax than the Fauvel and lower surface forced transition to smooth out the Cm curve doesn't affect their lift like it does the Fauvel.
Thanks Norman that's great.

I put those 3 into this comparison with mine and did the 0, 1 and 80% Bot setting for Transitions.

The affect of straightening out the CM between 2 and 12 Degrees is noted and is also present with mine.

The NM 69 has a good peak of CL and LD at the 9 Degrees AOA range, CLIMBOUT.

I dont know what to do about which Transition settings to standardize on, because they ultimately push the airfoil development into one direction or the other...

When I get some more time I will swap over to my CFD tool and see if it can help with comparisons/choices of settings.
 

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WINGITIS

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Ok Folks

Here is the CFD comparison to XFLR5.

I have used the NACA M6 airfoil as its a most straight forward one that everyone can see.

I had to close the TE in as the CFD does not like the open rear(Its attached)

For Reynolds 3,000,000 that represents a 1.3 meter wing in the 8 Meter CFD Tunnel at 0.1 MACH.

At the 17 Degrees tested it shows 1.25 Meters as the length but that is because it is angled.

It took 12.5 minutes to process the CFD analysis, I give it a lot of time to zero in on a quality answer.

RESULTS

1: Transitions NOT set, IE 0 Top and 0 Bot are closer to the CFD reported CL value, see CL screen of XFLR5.

2: The CFD shows a force line with angle at 0.357 Chord, which appears to be more in line with reality than the XFLR5 Arrow at 0.216 Chord that is always straight up.

NOTE: This is just a single comparison test and does not represent a definitive answer, but it would tend to indicate that we should not be using
Forced Transitions in our analysis/testing.

Does anyone know what the vertical arrow in the XFLR5 Pressure screen op point is supposed to indicate?

Additionally I fired up an old copy of DesignFOIL and it also got a closer result to the unforced transition option in XFLR5, but there are a lot of settings in this tool that can affect the result so I am not sure as to the validity of that.

Thoughts everyone?
 

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