Short Skyvan/SD330/SD360 Lift Distribution

[Mechta]

Thank you to everyone who has replied so far.

Brian Abraham's link is very informative, although it is largely about bodies with a circular cross section. It would appear that all conventional fuselages will develop some degree of lift if given an angle of attack, the only question is, how much?

I guess that testing a Skyvan fuselage by itself at different angles of attack will give the results that I want. It is interesting that the elevator extends right across the back of the fuselage, thus giving some control over the fuselage pitching effect (could also be for simplicity of manufacture). This is the best shot I can find of the tail surfaces (look at the amount of down elevator and how little actual tailplane area there is):



I found also this, Longitudinal Static Stability but have not yet crunched numbers with respect to the Short Skyvan family: It does not take account of the cross-sectional shape of the fuselage though, which seems odd, as one would expect a circular section to generate a lot less lift than a square one.

We have now most of the pieces required to predict the airplane stability. The last, and important, factor is the fuselage contribution. The fuselage produces a pitching moment about the c.g. which depends on the angle of attack. It is influenced by the fuselage shape and interference of the wing on the local flow. Additionally, the fuselage affects the flow over the wing. Thus, the destabilizing effect of the fuselage depends on: Lf, the fuselage length, wf, the fuselage width, the wing sweep, aspect ratio, and location on the fuselage.
Gilruth (NACA TR711) developed an empirically-based method for estimating the effect of the fuselage:


where:
CLaw is the wing lift curve slope per radian
Lf is the fuselage length
wf is the maximum width of the fuselage
Kf is an empirical factor discussed in NACA TR711 and developed from an extensive test of wing-fuselage combinations in NACA TR540.

Kf is found to depend strongly on the position of the quarter chord of the wing root on the fuselage. In this form of the equation, the wing lift curve slope is expressed in rad-1 and Kf is given below. (Note that this is not the same as the method described in Perkins and Hage.) The data shown below were taken from TR540 and Aerodynamics of the Airplane by Schlichting and Truckenbrodt:
Position of 1/4 root chord
on body as fraction of body length (X)
X Kf
.1 .115
.2 .172
.3 .344
.4 .487
.5 .688

Any further comments or info much appreciated!