In an aircraft, roll, yaw and pitch are controlled by the use of ailerons, rudders and elevators. To roll the plane to the right or left, the ailerons are raised on one wing and lowered on the other, effectively altering the lift on both wings (in opposite directions). The now unbalanced lift, forces the plane to rotate(roll) around it center of gravity. The physics behind yaw and pith works in much the same way. My problem is calculating how a change in aileron position affects the lift of the wing (and thereby the torque). The same applies for rudders and elevators. I assume these properties vary from plane to plane. But despite looking through several data-sheets on various airfoils, i have not been able to find this information anywhere. All suggestions are very welcome.

Imagine the aileron (or other control surface) deflects 'down'. This increases the camber of the airfoil section and thus alters the angle of attack of the airfoil and its cross sectional shape (the airfoil becomes a little more curved and increases the angle between the chord line and the air stream).

Ignoring the latter affect, the lift generated by a wing is a nonlinear function of the angle of attack. It's generally linear at small values of this angle but as the angle increases the lift first increases to a maximum and then actually decreases as the air flow starts to separate from the wing surface. This is called the stall and so at this point, the airfoil is partially stalled. The speed of the air stream also affects at what angle this occurs. At high enough angle of attack (/airspeed) the flow separates completely from the wing and the stall is complete.

On the flip side, deflecting the aileron up causes a decrease in camber and angle of attack, thus decreasing lift.

To implement this in a simulation, simply limit the maximum deflection of the aileron or other control suface. For wings and the horizontal tail, simply increase the lift force linearly with deflection angle 'down' and decrease is linearly (but at a higher rate) on the 'up' side. This difference in rate is due to the fact that a small deflection up washes off more lift than the equivalent deflection down creates. Alternatively, ignore this affect and assume that the affects are equal. In actual aircraft the down aileron deflects further than the up aileron so that the affects are balance and the change in loads on the wings are equal.

As for the rudder, just increase the force in the opposite direction to the deflection and make it proportional to the deflection.

I hope this helps!

Cheers,

Timkin

[edited by - Timkin on July 23, 2002 10:55:31 PM]

Thanks for your time Timkin (again).

Im afraid i already knew most of what you describe though.

What i didnt know is that the lift is a linear(approximately?) function of aileron deflection. This means that i will only have to find the *maximum* lift generated by the control surfaces, and not a curve describing every possible deflection level.

Im still at a loss when its comes to actually calculating these maximum lift differentials.

I want to add, that i have never worked with anything like this before, so maybe im approaching it the wrong way. If anybody has experience simulating these things please speak up .

[edited by - newdeal on July 23, 2002 11:41:35 PM]

Hello NewDeal,

(patronizing recap - I''m assuming you know this but..)

As Timkin suggests a wings lift coefficients are approximately a linear function of angle of attack until near [wing] stall. After which they ''drop off''. - I believe this is true for most ''normal'' aerofoil sections in sub-sonic conditions.

For a given aerofoil section you should be able to obtain more exact data on: how linear and where stall occurs - search the web or suitable text book.

Once you''ve found a lift coefficient this combined with the relative air velocity will produce lift and (a component of) drag.

(end patronizing bit)


Hopefully this should be the useful bit.

An alternative way you might choose to model an aileron, wing combination is as two wings.

----- . ---
(main wing) (aileron wing)

When the aileron is deflected this ''second'' wing has a additional (+,-) angle of attack.

This has (I guess higher computation cost?) But is more realistic allowing each ''wing'' to stall separately.

This method has the advantage of re-using code and equations you need to develop for a ''simple'' wing model.

Alternatively this method could be used to find the limits of the simpler model outlined by previously by Timkin. (How far can you push the aileron before it all goes horribly wrong)

Further realism can be added by introducing ''interference'' effects such as the second wing seeing an adjusted angle of attack due to first wing. (The presence of the forward wing lessens the second wings angle of attack)

Hope this helps

Tom.

quote:

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Original post by NewDeal
Im still at a loss when its comes to actually calculating these maximum lift differentials.

---

Okay, so assume you have the lift at zero deflection for the given flight condition. Then what you need is a lookup table that gives coefficient of lift and coefficient of drag for the given angle of attack. You can get this data off the web. Try the NASG airfoil database or one of the many other online airfoil databases (a google search using ''airfoil database'' should keep you busy for hours!).

Then, for the given deflection of the control surface, compute the change in angle of attack and determine the change in lift and drag coefficients. Plug these new values into your lift and drag equations and voila!


I hope this is what you were looking for!

Cheers,

Timkin

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