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Horisontal Stabiliser Angle of Incidence

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Battson

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Hey all,

Been such a long time since my days of study that I can't remember how to do this, can you help:

I am putting ribs on the tail of my Bearhawk, to change the horisontal stabiliser's shape from a flat Maule-style stabiliser, to a symetrical aerofoil shape. This is a standard "upgrade option" offered by Bearhawk Aircraft, except I've made my own ribs from a drawing they provided.

Now, obviously the aerofoil is more efficient than the flat stabiliser, so the native angle of incidence built into the tail has to be reduced (which is the whole point - grab a few kts for free, improve control effectiveness). The guys in the community have only recently started installing this upgrade, and I am not aware of any flying example (yet). Certainly there's no "bank" of data to compare and contrast.

Is there a straightforward and relatively accurate way of calculating (estimating) what the reduction on incidence should be?

The number being banded around the support group is about 50% reduction (from 4o to 2o), but I am interested to check the math and learn something in the process.


I realise some people find the Bearhawk "twitchy" in the pitch axis, so I'll say now that increasing control sensitivity is not my goal. Mostly I am after reduced drag for airspeed, and to a lesser extent, increased effectiveness for low speed ops with CG near the limits.

All thoughts welcome.
 

Head in the clouds

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The certified Drifted had a flat tailplane section and one well versed chap started to put taped foam ribs in his to improve the effectiveness of them. I did mine when I tried his, it had a much more positive feel. Nothing changed in terms of the pre-sets unless the CG was out, in which case the increased effectiveness resulted in less trim offset in level flight at cruise speed.

If your fixed stab is adjustable at the front or rear mounting I'd just set it as the plan suggests, fly it with mid range CG, note the trim position and adjust it accordingly.
 

Jan Carlsson

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Will there be a big change in CL/alpha slope going from flat to airfoil shape? is the new airfoil symetrical? I don't think there will, what is the alpha from start?

That 4-2 deg is that negative or positive angle?, the alpha depends on downwash from wing and speed/angle of fuselage.

Cheking in JavaFoil
a flat 4% thick round LE -flat - sharp TE CL/alpha 0,072
NACA 0009 CL/alpha 0,092 The same for NACA 63-64-009 series
 
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Dana

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dCL/dAlpha is pretty much independent of airfoil shape, so if you're going from a symmetrical section to another symmetrical section it shouldn't change the required incidence. If you're using a different non symmetrical section, then it would change as the zero lift angle would change and the the Cl/Alpha curve would shift over.

-Dana

What do you do when you see an endangered animal eating an endangered plant?
 

Battson

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I guess the best way to think of it, is going from a flat board (poor lift producing qualities) to a symetrical aerofoil (which is more effective at producing lift). There would be no difference at 0 incidence - but we have 4 degrees "angle of attack" in this case.

The angle on incidence is 4 degrees in the negative sense with the original design. The designer has said the change will result in a reduction of the angle on incidence of some kind.
 

wsimpso1

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Well, the dCl/dalpha of most foils is pretty much the same, even the tube structure based ones, so I would not think that the ideal incidence of either should be different from the other.

Remember also that we usually use a symetric foil for tails because in cruise, the tail is usually running at pretty small Cl. The big deal is to set the tail to minimize drag in cruise. If ever there was an excuse for running a little of that sublimating oil it would be to go do a cruise test on that model airplane with either tail and find out what the actual cruise AOA on the tail is.

When in doubt, do what the designer says.

Billski
 

Dana

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I guess the best way to think of it, is going from a flat board (poor lift producing qualities) to a symetrical aerofoil (which is more effective at producing lift). There would be no difference at 0 incidence - but we have 4 degrees "angle of attack" in this case.

The angle on incidence is 4 degrees in the negative sense with the original design. The designer has said the change will result in a reduction of the angle on incidence of some kind.
No, it shouldn't matter. A good airfoil isn't "more effective at producing lift", but it's more efficient, i.e. the L/D is better. The lift produced at a given AOA is the same, but the flat plate will have more drag. Also Clmax will be lower (and reached at a lower AOA) for the flat plate.

-Dana

I think we've solved the energy crisis. Just hook a turbine to the graves of the Founders. Washington, Madison, Adams--they must all be spinning in their graves. Rapidly.
 

Head in the clouds

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No, it shouldn't matter. A good airfoil isn't "more effective at producing lift", but it's more efficient, i.e. the L/D is better. The lift produced at a given AOA is the same, but the flat plate will have more drag. Also Clmax will be lower (and reached at a lower AOA) for the flat plate.
Yes, with the example I gave of the Drifter below the overall effect was minimal except perhaps the drag was lower but in a Drifter that would be hard to notice. The main noticeable effect was with very little elevator deflection, the flat plate seemed to do nothing until a degree or two of deflection, whereas the shaped foil didn't have that apparent 'dead spot'. Bearing in mind that the Drifter tailplane is in a mess of turbulent wash coming off the unfaired fuselage.
 

Battson

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No, it shouldn't matter. A good airfoil isn't "more effective at producing lift", but it's more efficient, i.e. the L/D is better. The lift produced at a given AOA is the same, but the flat plate will have more drag. Also Clmax will be lower (and reached at a lower AOA) for the flat plate.
You must be assuming the lift is only generated by the angle deflecting air flow, in that case? A good aerofoil like an asymetrical wing produces lift *because* of it's shape, so saying it's independant of the aerofoil shape doesn't make sense. Whether it still applies for a symetrical aerofoil, I dont know.
The induced drag is a function of that. I'd have called it form drag in the case of a planar surface at an angle to the airflow, because Bernoulli's theory doesnt cause the force - it's just the momentum change.
 

Dana

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You must be assuming the lift is only generated by the angle deflecting air flow, in that case? A good aerofoil like an asymetrical wing produces lift *because* of it's shape, so saying it's independant of the aerofoil shape doesn't make sense. Whether it still applies for a symetrical aerofoil, I dont know.
The induced drag is a function of that. I'd have called it form drag in the case of a planar surface at an angle to the airflow, because Bernoulli's theory doesnt cause the force - it's just the momentum change.
We were talking about symmetrical airfoils. The lift curve slope for an asymmetrical airfoil is the same, but it moves over... the zero lift angle is negative, so the entire curve moves up. We speak of the absolute AOA, i.e. measured from zero at the zero lift AOA, and from that point airfoils behave the same, up to where lift falls off at it approaches Clmax (stall).

-Dana

I think we've solved the energy crisis. Just hook a turbine to the graves of the Founders. Washington, Madison, Adams--they must all be spinning in their graves. Rapidly.
 

Battson

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Ok, yes we were.

So I am to take it that there should be no reduction in AoA for the tailplane, given a symetrical aerofoil.
I wonder where the community got their 2 degree reduction figure from..........
 

Autodidact

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Here's a quote from a paper I found on Purdue's website:

The effect of thickening the flat plate is to increase the lift curve slope dCl/dalpha slightly as can be seen from Figure 11-1 and Figure 11-2. However, this theoretical prediction is not observed experimentally, probably because of the viscous effects that are neglected in the inviscid theory. With respect to force and moment, the flat plate can be considered as a limiting case of a symmetric airfoil, as the ratio of thickness to chord approaches zero. Flow about a Circular Arc Airfoil is shown in Figure 11-3.
Here is the link:

http://www.google.com/url?sa=t&rct=j&q=&esrc=s&source=web&cd=7&ved=0CFYQFjAG&url=http://www2.tech.purdue.edu/met/courses/met490/MET490airfoillab.DOC&ei=FvNsUIuvHtSTqwGA1oGgAQ&usg=AFQjCNEHGh0arAgIAbWtSeozWv-csqgwuQ

You might want to talk directly to the designer; things can easily be misunderstood and then the misunderstanding propagated as the info passes from one person to another...
 

Battson

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That is interesting.

My research turned up that people have done tests with the flat "aerofoil" tailplane for the Bearhawk, and they found that the aircraft wont stall once the angle is shallower than 2o, as the angle approaches zero the aircraft becomes twitchy and exposed to trim tab flutter. In the end they returned the orginally designed angle, even though it does not have the elevator "trailing" straight behind in the lowest drag configuration.

I speak to the designer one in a while, so I will ask directly when next I do.
 

wsimpso1

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I have always found folks commenting on dCl/dalpha of different foils to be splitting hairs. When you first take the airfoil curve, then correct for finite aspect ratio and then consider the chopped up flow from the prop and the fuselage and leakage through the gap between elevator and stabilizer, well, any little difference between foils becomes moot. In reality, all foils have the same dCl/dalpha slope. Dana understands this - listen to him.

In addition, all foil theory says that the form of the foil is irrelevant to the lift curve slope. While you may pooh-pooh theory, it turns out to have panned out in real foils. Read TOWS and believe...

Now why would the designer ended up with a non-optimal tailplane incidence? Bunches of reasons exist. First and foremost is that it will not likely make much difference if they are using a tube formed tailplane - they are draggy and very leaky to begin with, so optimizing the incidence is way down on the list of priorities. Next, if the elevator was unloaded at cruise, it may have had any number sins, such that a slightly non-optimal incidence plus a deflected trim tab gives a much nicer flying airplane.

Putting a better foil on the tail might take off some drag counts and add a little control responsiveness, but do not expect big changes in cruise speed, even if the incidence is perfect. The tailplane is just not that big a contributor. If you want the airplane to be faster, you really need to do a comprehensive drag cleanup. Gap seals, root fairings, engine cooling, fairing hardware and the landing gear, and then optimizing the propellor all contribute.

Billski
 

Jan Carlsson

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That is interesting.

My research turned up that people have done tests with the flat "aerofoil" tailplane for the Bearhawk, and they found that the aircraft wont stall once the angle is shallower than 2o, as the angle approaches zero the aircraft becomes twitchy and exposed to trim tab flutter. In the end they returned the orginally designed angle, even though it does not have the elevator "trailing" straight behind in the lowest drag configuration.

I speak to the designer one in a while, so I will ask directly when next I do.
This is probably the reason, an airfoil on the tail, can or will, have a higher CL max, due to very slightly steeper CL/alpha slope, that can make it possible to have smaller angle on the tail, the tail incident is not only sat for minimum drag, it is often necessarily to help the elevator to flare in ground effect with flaps down. at a 3-pont landing the alpha of the tail plane is 10-15 degree with elevator up, then 2 degree is 13-20%


If Java foil is correct ,092/,072= 28% difference, even if the numbers isn't 100% correct the difference will be in the same order.


Normally as Billski point out the difference from airfoil to airfoil is veeeery small. Also confirmed by JavaFoil, but it indicate a significant diff when it come to a flat /tube round nose airfoil.
 

Autodidact

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Here's a picture of a Taylor Monoplane; you can see that there is quite a bit of incidence on the stabilizer, even in the face of some down-wash from the wing:
 

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Battson

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This is probably the reason, an airfoil on the tail, can or will, have a higher CL max, due to very slightly steeper CL/alpha slope, that can make it possible to have smaller angle on the tail, the tail incident is not only sat for minimum drag, it is often necessarily to help the elevator to flare in ground effect with flaps down. at a 3-pont landing the alpha of the tail plane is 10-15 degree with elevator up, then 2 degree is 13-20%


If Java foil is correct ,092/,072= 28% difference, even if the numbers isn't 100% correct the difference will be in the same order.


Normally as Billski point out the difference from airfoil to airfoil is veeeery small. Also confirmed by JavaFoil, but it indicate a significant diff when it come to a flat /tube round nose airfoil.

I am not sure I understood what the percentage values are - is that the change in alpha valve with a reduced horisontal stabiliser angle?

As you mention, the Bearhawk tailplane is normally a round tube leading edge with a flat surface back to the trailing edge. The stabiliser has parallel top and bottom surfaces, but the elevator tapers to a smaller trailing edge. The elevator on the Bearhawk is much larger than the stabiliser, even without considering the horn balance.

BH3viewa.jpg
 
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clanon

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and they found that the aircraft wont stall once the angle is shallower than 2o, as the angle approaches zero the aircraft becomes twitchy and exposed to trim tab flutter.
Not enough Longitudinal dihedral... maybe... by the drawing is about 6 degrees
(2deg Wi + 4deg HS)
 

clanon

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The basic would be:
It will make more CL (how much...)
Then it will need a lower (closer to 0) angle
What airfoil is it 0012 0009...?
MAC ?
RN or speed ?

PS:please feel free to correct me.
 
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