Discussion Thread: The design of a tailless flying wing

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Aerowerx

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I am not as concerned about BSLD as I am about proverse yaw. IMHO having a specific shape distribution is not as important as the overall features. Such as proverse yaw and max CL at 30-40% of half span.

I used Nickel's formula as a starting point, then tweaked them in XFLR5 until I got no, or very small (less than 1 degree), pitch change with flaps deployed.
 

pictsidhe

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Nickels flapping formulae aren't super accurate anyway. Good enough for preliminary spreadsheet.
I doubt many aircraft really have an elliptical distribution, reality and practicality intervenes and they end with something sort of ellipticalish. You have to stray quite a way for it to make a measurable difference. With your centre lift, you'll have a bellish distribution.
 

pictsidhe

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The advantage of specifying some function as the distribution comes in the related math. Arbitrary ones are not friendly to analysis.
 

Aerowerx

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N1M: (Discussion Thread: The design of a tailless flying wing)

I was looking at the Northrop N1M and N9M, to see what airfoils they used.

I see that the N1M uses use a symmetrical airfoil for the root section with 25% thickness. In order to have the cockpit entirely in the wing you have to use these thick airfoils.

Just to see how it worked I tried my XFLR5 model with a symmetrical airfoil (NACA 0012). Interesting. I get the same general shape of the lift distribution as with the low-moment cambered airfoils.

Anyway...

In this old thread (post 33), Norman attached a diagram of the N1M. Here is a snip of the portion of it pertinant to my question.

Capture.JPG

It appears that the pilot is sitting in a rather cramped position with rather poor visibility. Is this correct? I think it would be rather uncomfortable on a long flight.
 

Norman

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Re: N1M: (Discussion Thread: The design of a tailless flying wing)

I see that the N1M uses use a symmetrical airfoil for the root section with 25% thickness.
The N1m, N9m, and XB-35 all used the same 19% airfoil at the root and an 18% version of the same airfoil at the tip. It was an entirely inappropriate choice for the little N1M because it had massive separation at the low Reynolds numbers. It wasn't a good choice for the full size bomber either because the thickness limited top speed and the CLmax of symmetrical airfoils is low which limited the altitude that it could cruise at with a full load. Cambered airfoils are more efficient at higher CL than symmetrical ones.

In order to have the cockpit entirely in the wing you have to use these thick airfoils.
Only a 2 or 3 foot wide area root. The rest of the wing should be thinner. Anything thicker than 17% starts losing CLmax and control surfaces are most effective on airfoils 13% or less depending on the local Reynolds number.

Just to see how it worked I tried my XFLR5 model with a symmetrical airfoil (NACA 0012). Interesting. I get the same general shape of the lift distribution as with the low-moment cambered airfoils.
Why should it be surprising that the lift distribution didn't change very much? All airfoils have the same slope to the cl over alpha curve so changing the profile shouldn't have much affect on the lift distribution.

It appears that the pilot is sitting in a rather cramped position with rather poor visibility. Is this correct? I think it would be rather uncomfortable on a long flight.
They considered it a sub-scale mockup (that's what the M in N1M stands for) for testing wing and control surface configurations. Pilot comfort wasn't a consideration. It didn't have the power to actually go anywhere so flights usually weren't very long. When they moved it from the factory in Hawthorn to the Muroc dry lake they towed it with a C-47 because it couldn't cross the Sierra Nevada under its own power.
 
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Aerowerx

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Re: N1M: (Discussion Thread: The design of a tailless flying wing)

Only a 2 or 3 foot wide area root. The rest of the wing should be thinner. Anything thicker than 17% starts losing CLmax and control surfaces are most effective on airfoils 13% or less depending on the local Reynolds number.
I was wondering about that, and it is on my list of things to investigate in my virtual wind tunner.


Why should it be surprising that the lift distribution didn't change very much? All airfoils have the same slope to the cl over alpha curve so changing the profile shouldn't have much affect on the lift distribution.
But the angle at zero lift is different for each airfoil. Would that have an effect on the lift distribution? The PRANDTL-D, and my previous models, has different airfoils at the root and tip, and morphs linearly between them. In this model I used the same airfoil on the entire span.

Besides, isn't it the distribution of lift coefficient that is important and not the actual lift distribution? You want it to be negative at the tips and maximum at about 1/3rd semispan.
 

Norman

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Re: N1M: (Discussion Thread: The design of a tailless flying wing)

But the angle at zero lift is different for each airfoil. Would that have an effect on the lift distribution? The PRANDTL-D, and my previous models, has different airfoils at the root and tip, and morphs linearly between them. In this model I used the same airfoil on the entire span.
Yes, Cm0 and ZLA are factors in calculating washout but since the airfoils have the same lift slope and aerodynamic washout is relative to the ZLA the lift distribution should be pretty similar. Since, I assume, you're using the twist values that you calculated for the Prandtl-D airfoils and simply replacing them with the symmetrical airfoil I would expect the amount of total twist to be wrong and the plane will trim at a different airspeed. Since the Prandtl-D airfoil has a small negative pitching moment it needs more washout to trim than a symmetrical section. The result is that when you substituted the symmetrical airfoil cross section the new model had too much washout and that would make the tip CL go negative at the same cruising speed as the first model.

Besides, isn't it the distribution of lift coefficient that is important and not the actual lift distribution? You want it to be negative at the tips and maximum at about 1/3rd semispan.
The tips should only go negative if the plane is flying faster than the speed it is trimmed for i.e. when it is accelerating in a dive with controls fixed.
 

Aerowerx

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The design of a tailless flying wing: Pushers and self-trimming flaps

Most tailless aircraft are pushers.

On a self-trimming flap, it seems to work best when the in-board end of the flap is at the wing center. Now this is not entirely possible, because you need room for the pod/fuselage, not to mention the engine.

So the configuration I have been looking at has the in-board end of the flap at 1 or 2 feet out from the wing center line, as seen here...
Capture.JPG
This is a right-rear quarter view.

And perhaps more clear in a right front quarter view...
Capture.JPG

Although the engine thrust line will be above the wing, I am still concerned about having the flaps in front of the prop. Although I haven't got that far yet, I have been thinking about a 4 or 4.5 foot 3-blade prop, so there may only be 4 or 4.5 inches of the flap in front of the prop. The prop will also be about 2 feet behind the wing.

Is this still a concern or do you think it will be ok? Other than maybe increased noise with flaps deployed, would there be any other problems?
 

Norman

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As long as the most efficient area of the prop (about 75% of the radius) is well away from the wing's wake flaps shouldn't cause a big problem as far as thrust is concerned. The downwash extends a couple chord lengths above and below the wing. The downwash angle and velocity profile entering the prop will change some with flaps down but since it's clean air, except for the few inches in the wake, the prop won't be affected very much. However the prop has a turning effect that will pull air away from the section of the flap directly below it and probably increase separation there. Not a big deal, just a bit more drag from the flap
 

Aerowerx

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All wings fly, but only the poorly designed ones need a tail.;)

If you look at my OP when I started this thread you will see that I allowed for fins.
 

skysoarer

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Could someone help me with my understanding. Here is a quote from the conclusion in a paper by A Bowers.
" Neither Prandtl nor Horten followed through to the logical and complete conclusion of their work. Prandtl did not extend the upwash outboard of the wingtip, which would have answered the question of formation flight in birds, and he did not find the induced thrust at the outboard ends of the wings, which leads to proverse yaw. In turn, with his approximation and objectives Horten did not understand the origin of the induced thrust at the outboard ends of the wings for proverse yaw, and he did not prove that proverse yaw exists.
It remained for the current authors to prove conclusively that proverse yaw is achievable through an efficient bell-shaped spanload, that an optimal solution integrating minimum structure and minimum drag can solve the problem of yaw control and stability of a flying wing, and that the bell spanload solution answers some of the great enduring mysteries of the flight of birds.
In the case of the flight of birds, the bell spanload is the only viable solution."
End of quote.
Would I be right in thinking that twist in the bell spanload distribution is the cause ( as well as wing shape) of proverse yaw and in which case, why are not wingtips deformed downwards at speed as has been the case with Horten gliders and other aircraft.
Birds only have lightweight structures with which to contain these forces and yes, I appreciate they can manipulate there planform, but if this bell spanload distribution is correct and it does seem to make sense, the author ( A. Bowers) mentions the wingtip feathers are not deformed, so can someone please put this in a simple context that I can grasp.
Many thanks
 

Aerowerx

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...
Would I be right in thinking that twist in the bell spanload distribution is the cause ( as well as wing shape) of proverse yaw and in which case, why are not wingtips deformed downwards at speed as has been the case with Horten gliders and other aircraft.
Birds only have lightweight structures with which to contain these forces and yes, I appreciate they can manipulate there planform, but if this bell spanload distribution is correct and it does seem to make sense, the author ( A. Bowers) mentions the wingtip feathers are not deformed, so can someone please put this in a simple context that I can grasp.
Many thanks
I'm not sure what you are asking.

What do you mean by "deformed downwards at speed"? The tips are moveable?

To get the proverse yaw, which is caused by induced thrust there has to be a washout at the tips. If you have the PRANDTL-D paper, there is a table towards the end with the wing twist used. You will note that the tips are twisted so the leading edge is pointed downwards. You will actually get a negative coefficient of lift at the tip. This can be seen easily in XFLR5, but setting up the nonlinear twist is quite tedious.

It is hard to see with birds in flight, but the few times I have been able to see it, the tip feathers are angled in the same way.
 

Norman

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Would I be right in thinking that twist in the bell spanload distribution is the cause ( as well as wing shape) of proverse yaw and in which case, why are not wingtips deformed downwards at speed as has been the case with Horten gliders and other aircraft.
Birds only have lightweight structures with which to contain these forces and yes, I appreciate they can manipulate there planform, but if this bell spanload distribution is correct and it does seem to make sense, the author ( A. Bowers) mentions the wingtip feathers are not deformed, so can someone please put this in a simple context that I can grasp.
Many thanks
Yes, it's the washout that makes this work. Sweep and taper both cause the circulation to be weakened at the forward end of a swept wing and strengthened at the other end. This results in a shallower slope of the cl over alpha curve at the root of a swept back wing and a steeper curve near the tip. This steeper curve of the lift slope is what causes swept wings to develop a pitching moment in addition to the 2D airfoil pitching moment and tip stall. A small degree of washout can fix the pitching moment problem but only partially alleviates the tip stall problem, thus swept wings usually have some sort of stall delaying device near the tips. The trailing vortex (often called the tip vortex because that's where it usually starts) is part of the wings circulation. Looking at the flow field fro behind you can see that the inboard side of the vortex is going down and the outboard part is going up. The inboard part of the vortex adds to the downwash behind the wing and the ascending part outboard is normally lost. By forcing the vortex to start inboard of the tip a small part of the span (about 22%) is in the upwash side of the circulation. This little piece of wing outboard of the vortex is like a glider flying in slope lift. The lift vector is tilted forward just as a glider diving into an updraft. When you increase the camber of the wingtip the lift there increases but since the lift vector is tilted forward the induced drag on that part of the wing does not increase.

why are not wingtips deformed downwards at speed as has been the case with Horten gliders and other aircraft.
If you force the plane to fly a lot faster than it was designed for the wingtips will start lifting downward. With a swept flying wing the only way to exceed the trim speed is in a sustained dive and even then the plane will pull out by itself because the tips are your pitch control surfaces.
 
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