I had at the back of my mind something that I had seen about this stuff and I finally remembered it. Go to

http://www.desktopaero.com/appliedaero/appliedaero.html

There are no equations in this page, very basic assumptions - No fuselage, no wing section pitching moment, just basic stuff.

The definitions of the inputs in the figure are below

sm = static margin, the measure of stability based on the reference chord. Distance between the CG and the Neutral Point.

Wing AR and Tail AR = changes in the picture as you change the numbers. What you see is what you get

St/SW = the ratio of tail area to wing area, easy stuff.

The Figure changes to show the changes, you can click/drag the tail left and right and see how the changes occur.

If the Lt/Lw printed out is positive it indicates a tail up load.
e is the induced drag and is not too interesting.

I chose the following for an example of a aerobatic ship.

sm = .10
Wing AR = 6
Tail AR = 3
St/Sw = .25

Push the compute button.

Use the mouse and move the red horizontal tail left and right and see the load on the horizontal change. If the term Lt/Lw is positive then the load is up.

For the fun of it try a sm = .4 (really stable) and see the negative values appear.
Put in a sm = .25 and see the load change about half way back as you move.

It is a good tool to find out what is happening. Of course if the airfoil has a big moment input or the fuselage, flaps, etc it will change the answer. Sometimes not easy to guess at.


---------- Heaven help me.... Now I'm quoting my own stuff.... Looking around for items relating to center of pressure (totally forgotten during this morning's very early 1:30 AM ephiphany) I'm seening lots of conflicting stuff. It seems that the old way of looking at an overall aircraft center of pressure has been tossed out and this new Neutral Point method put into it's place. --------------

Well I argue with myself and I always lose or win depending on which of me wins. It gets worse late at night. My wife mentions something about looney toons as she walks by.


------------
The Neutral Point (NP) consideration calls for it to be fixed based on the geometry of the airplane. Similarly the CG is fixed. But something has to move or the airplane would not take any corrective action when upset in flight.

This brings me to the overall aircraft lift Center of Pressure (lets call it the CPa). This being the resultant point of all lift forces of the wing, tail and fuselage and (YES), even drag moments. So .................. center of drag. In fact in the diagram I've shown below even though I've shown the CPa moving forward and up it's equally possible that it could move up and BACK as long as there was still a positive righting force from the drag to overcome a rearward shift of lift. Strange but true as I see it.----------

No, you don't need to approach the problem with a CPa thing. The answers will fall out without it. Assume constant velocity.

When you are neutrally stable the upset motion (say increasing angle of attack) can be made without any restoring moment from the tail.

If you have a stable airplane the moment due to the stab INCREASES FASTER than the moment due to the wing. Thus you get the restoring moment.

How it works is for a neutrally stable airplane is this

Cmwing = Cmtail

Cmwing = (Clalphawing X alphawing X wingarea X dynamic pressure) X distancetoCGwing
Cmtail = (Clalphatail X alphatail X tailarea X dynamic pressure) X distancetoCGtail

crossing out things that can be equal we get

wingarea X distanceCGwing = tailarea X distanceCGtail

But if the airplane is stable the distanceCGwing number has decreased from the neutrally stable point and the distanceCGtail increased from that point. This means that if disturbed the tail moment will increase faster than the wing moment. It works for pylon free flight or Extras equally well.

If the velocity is increased and the airplane is stable then the fact that the tail moment increases faster than the wing moment gives a nose down and a possible tuck under if the wing incidence angle is close to the tail angle. To stop this make the wing incidence angle greater than the tail angle to get angle of attack stability, put in down thrust to counter a pitch up in the climb which is faster than the glide speed and you get an airplane that works. Of course the spiraling climb works better.

Of course if you have a pylon model then those effects have to be considered as the two drag differences, wing and tail, will make a moment also. But it is a moment thing and not a moving center of pressure. You would have to determine the vertical center of gravity and see if the drag effect moment arms work out.

I have never done it so don't know how it would look.

I try to keep the concepts of pitch stability and pitch trim in two seperate compartments of my thinking process.

The neutral point is established by the plan of the aircraft and the effectiveness of the flow over the wing and tail. Consider it essentially fixed (with minor exceptions) after the aircraft is built. The static stability is established by where the CG is placed relative to the neutral point. There may be some change in CG location between a full tank and an empty tank. This is accomodated by making the static margin large enough so that, with the CG in the aft most location, the stability is adequate.

The pitch trim is established by balancing all the forces and moments. This is done to establish a trimmed airspeed by adjusting relative angles between the thrust line, the chord line of the wing's mean aerodynamic chord and the horizontal tail chord line. The thrust line is also adjusted for a minimum change in pitch attitude over the throttle range. To make this proceedure practical, two of the three critical angles must be adjustable. Notice that the word incidence isn't involved except possibly as an intermediary between the three functional and critical angular relationships.

By adjusting back and forth between stability and pitch trim the model can be brought near its ideal setup.

Now what's so complicated about that?

It is pretty easy after all although incidence is a perfectly good word. Might as well use it.

However a big light bulb lights up when someone says that tail lift is ALWAYS down. Just not the case.

Very interesting!
I guess this tread is more or less completed, but anyway....

I have seen some of you have been mentioning sweept flying wings, witch use wing-twist (washout) rear of NP, or airfoils with positive Cm to get a positive pitching moment around NP. CG is then located in front of NP to counteract and make stable flight.
If the twisted wing tips (lets say -2°) create downforce or upforce is not really relevant, this will vary upon angle of attack and the importent thing is that they provide a positive pitching moment around NP.

As I see it, this is not different from conventional planes, witch use the tail to get a positive pitching moment around the NP.
As mentioned before in this tread even a fully symetric acro model need a few clics of uptrim to fly stable. When inverted the same plane need a little down elevator to "turn" the pitching moment the other way around.

Isn't it then correct to say that the tail on that same plane totaly creates a lift when AOA is positive (level flight), but it also creates a downforce relative to the main wing due to the up-trim?

Of course if the mainwing airfoil has a negative moment the stab must have more negative AOA relevant to the mainwing to counteract so that the total pitching moment around NP is still positive.

In other words, whether the tail is creating lift or not change and isn't really relevant to stability, but the pitching moment it creates around NP is.

or ....?

Olav, Norway

Oleg,

In the non-accelerated flight (eg. level flight at a constant airspeed) of a stable aircraft, the sum of all the vertical forces is zero, the sum of all the horizontal forces is zero and the sum of all the moments is zero (about any point you choose). The plane adopts a pitch attitude that produces an angle of attack of the wing that balances the forces and moments at that airspeed. That angle of attack includes the induced angle of attack which is a function of aspect ratio and coefficient of lift. As the airspeed changes slowly to another airspeed and stabilizes at the new airspeed the pitch attitude of the plane changes to accomodate a new balance of forces. Across the full range of stable pitch attitudes associated with the unstalled angles of attack of the wing, the stabilizer will also adopt a range of angles of attack which contribute to the balance of forces and moments. Since the stab is operating in the downwash of the wing, the stab angle of attack is affected by the wing's downwash angle which increases as the square of the wing's coefficient of lift. The stabs range of angles of attack may carry it through both positive and negative lift conditions.

--------------------- As mentioned before in this tread even a fully symetric acro model need a few clics of uptrim to fly stable. When inverted the same plane need a little down elevator to "turn" the pitching moment the other way around.

Isn't it then correct to say that the tail on that same plane totaly creates a lift when AOA is positive (level flight), but it also creates a downforce relative to the main wing due to the up-trim? -------------------

With the fully symetric acro model the few clicks of up trip lessen the up force of the tail, not create a downforce relative to the main wing. Another way of saying it is that it creates an incremental load that is negative while the total load is still down. Up-trim is not the same as downforce, just an increment.

-------------- Of course if the mainwing airfoil has a negative moment the stab must have more negative AOA relevant to the mainwing to counteract so that the total pitching moment around NP is still positive. ------------------

If the negative moment of the mainwing airloil is big enough, or the CG is ahead of 25% chord, then the stab can end up with a down force. However the pitching moment is always zero around the NP. That is the definition of the NP.

---------------- In other words, whether the tail is creating lift or not change and isn't really relevant to stability, but the pitching moment it creates around NP is. --------------------

Since the moments from the tail comes from the lift of the tail (up or down) it is relevant. Where it becomes important is in designing structure. In full scale aircraft structure that is too strong for the loads is wasted. Looking back at the NACA report the engineers would look at the flight loads and perhaps redesign the horizontal tail or add bracing.

This one's still going? Wow.....

Just to add to my free flight study. I'd forgotten one little item of demonstable evidence that really proves that the contest free flights are flying with a very positive tail lift.

To add turn in the glide without using trim rudder, and thus affecting the climb, we tilt the stab. To get some right turn the stab is tilted right tip high left low. This points the upper surface off to the left but the model turns to the right. This only works because the stab is lifting positively and the tilted lift component is towards the left, just like some right turn tab.

If it was negative then this same tilt would have a component pulling to the right and causing a left turn.

I'd already figured out the other explanations but in case anyone out there needed a final bit of proof.....

Roger with the stab tilt. Interesting isn't it.

I thought a graph or two might help some with the direction and amount of tail loads involved in the typical symmetrical aerobatic model. Remember graphs are our friends. Don’t hold my feet to the fire with absolute numbers. I hope the plots are readable. I take no credit for mis types, mistakes and the like. This is just some representative numbers to show trends.

Definitions

dH is the horizontal tail deflection in degrees. The convention is that + is the deflection to cause an up load on the tail.
Alpha symbol is angle of attack in degrees.
CL is the lift coefficient. Up is positive.
CM is the pitching moment coefficient. Nose up is positive.

Think of CL as lift, Cm as the pitching moment.

These are curves that might come out of a wind tunnel. Looking at the CL verses angle of attack plot, for the zero tail deflection case, we see that the line goes through the 0-0 point of the axes. CL gets larger at high alphas and goes to large negative values at low alphas. This is what you expect.

The CL vs Cm curve for the zero tail deflection case goes through 0-0 and has for positive CL values a negative Cm (airplane nose down). This indicates that the airplane is stable and a small disturbance in angle of attack will return to the trimmed point.

Now for the tail deflections. Looking at the CL verses alpha plot for + values of alpha and + dH at each angle of attack the increment is up and the overall lift is positive as long as alpha is positive. When the dH is – the increment is down but the overall lift is still positive. Of course the trend is reverses at – alphas.

Looking at the CL vs Cm plot and for + values of dH the moments are negative which is airplane nose down. For – values of dH the moments are positive which is airplane nose up. Each line of dH data is parallel to the dH=0 line so the airplane remains stable but will go nose up or down as commanded.

The curves will shift as airfoil/wing and airplane-body lift and pitching moment characteristics change.

So what do the plots finally show?? An airplane flying in the upper right hand quadrant of the CL vs alpha plot will have tail loads that are up. The exception is for transients at the beginning of maneuvers at low angles of attack. Once the maneuver has obtained steady state conditions (with our models this is really rapid) the loads are up on the tail. Fly upside down and you are in the lower left quadrant.