Re: Wing Load Testing

Jason Kramb

Sorry....just catching up on this thread.

There's lots going on in the distribution of load between the wing and
canard, and it definitely depends on the CG location of the aircraft. It
also depends on the angle of attack, the altitude, temperature, and a host
of other factors.

Essentially, in a normal airplane, the wing-body (think of the plane with no
horizontal tail), has a certain amount of lift, which can be thought of as
acting on the center of pressure of the wing. It also has an inherent
pitching moment (normally nose down) which comes from the airfoil design.
The combination of that pitching moment and the weight of the aircraft
(acting on the CG, which is normally in front of the Center of Pressure, and
therefore also causing a nose down pitching moment), are balanced by the
downforce created by the horizontal tail.

In a canard, the same is true, except that because its in front, the canard
needs to produce lift to balance the nose down pitching moment. How much of
that pitching moment is due to the airfoil, and how much is due to the
distance between the CG and center of pressure of the wing determines how
much load ends on each wing.

Now finding the exact center of pressure is relatively hard to do because it
does change with angle of attack as does the pressure distribution on the
airfoil surface. What is assumed in aerodynamic design is that everything
is measured from the 0.25% chord location, including the inherent pitching
moment of the airfoil. For a 3D wing, that is normally translated into the
MAC (mean aerodynamic chord) and the quarter chord of that is used.

Jay is correct that the pitching moment from the wing airfoil will increase
as speed goes up, but then so will the lift generated by the canard to
balance it out. The neutral point of the aircraft (the point about which
the aircraft is neither stable or unstable) can be calculated using all
those airfoil parameters, as well as the location of the wing and canard,
and it doesn't move much until you start hitting transonic effects (way out
of our range). I would guess that the pitching moments on the airfoils are
small relative to the moment created by the distance between the wing and
the CG location. Its probably within your error margin to assume the lift
comes from the 0.25% MAC line on the canard and main wing and determine the
distribution based on the CG location. Certainly, guessing that its at
0.25% instead of 30% will give you a slightly higher load for an aft CG
location and be the more conservative answer.

I'd bet that since we tend to make a calculation based on a conservative
guess and not spending the time to get the exact answer, that's probably
what Burt did originally. However, back engineering something when you've
got very little of the original data is a lot harder than starting from
scratch. =) I'll have to ask Burt if he still has any of his design data
lying around.

As for where to put weight on a wing to test it, you are testing the spar,
so you would like to put the weight on the spar. Airfoil pitching moments
act on the entire wing and so don't twist the spar at all. Aileron
deflections can twist the wing somewhat if the wing is long and slender and
doesn't have ailerons the entire length of the trailing edge. For most GA
planes, this is not even an issue because the torsional stiffness of the
wing is pretty high. For the solid wing contruction of a Quickie, its even
less of an issue. So, the load should be applied as much on the spar cap as
possible, and in a spanwise distribution that is most representative of the
actual load. This is not uniformly from end to end, but an elliptical
shape. The exact shape is hard to determine without some detailed
calculations, but a standard ellipse is also probably within your error

Jason Kramb
Aero Design Engineer
Scaled Composites

On Sun, Jan 11, 2009 at 9:33 AM, JAY SCHEEVEL <> wrote:

Hi Guys,

I think you may be splitting hairs here. Jason Kramb, please correct me if
am wrong:

I believe that the percentage of load carried on each wing is a function of
airspeed. This is because of the pitching moment contribution of each wing.
The influence of pitching moment is significant and has been neglected in
discussion so far. Both airfoils (wing and canard) are assymetric (top to
bottom), they have trailing edge upward pitching moments that varies as a
function of angle of attack and airspeed.

Because of the contribution of pitching moment, the "effective" load
progressively to the canard with increasing airspeed. This means that at
lowest airspeeds, the highest load percenteage is on the rear wing
around 35%) and at the highest airspeeds, the highest load is on the canard
have heard some on the Q-performance list say it is as high as 90%, but I
not verified..I think this number comes from X=plane). So I think the
to pin an exact percentage on each wing the is a fruitless exercise.

As far as the chord at which you find the "center of lift", again this is
a number that is significant if there is no pitching moment contribution.
Since the wing is likely to fail by fiber crushing ion the top surface (not
torsion), it probably does not matter where on the cord the loads are
as long as they don't fall off during the experiment. To simulate a
edge up pitching moment, assuming that the wing is inverted for the
experiment, you would want to bias the weight toward the trailing edge as
load it.

Jay Scheevel -- Tri-Q still building


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