How to Calculate AIRCRAFT Lift & Drag Coefficients?

[green66]

Hi All,

Need some advice on how to determine aircraft lift and drag coefficients (without physical testing, that is).

My application is design of a thermal glider, where I want to get some idea of sink and lift / drag ratio (glide slope) in order to assess various combinations of wing planform, airfoil, aspect ratio, and wing loading.

I wrote a spreadsheet (see pic) to estimate sink rate and L/D, where the airfoil polar data is pasted from a Profili / XFoil airfoil analysis. I assumed aircraft lift coefficient equal to wing lift coefficient, and total fuselage + tail drag coefficient assumed constant at 0.01

The results seem realistic, and I believe the basic equations are correct, but I'm wondering if this simple non-CFD approach is a valid way to estimate performance? The spreadsheet is based on one Reynolds number, and doesn't consider wing planform or size. For a real wing, i.e. one with finite span and chord / taper breaks, the Reynolds number, lift force, and lift coefficient all vary with span location.

If the planform being analyzed provides a reasonably elliptical lift distribution, as predicted by a vortex-lattice based program such as John Hazel's LiftRoll, can the spreadsheet coefficients be factored or manipulated in a straightforward way to account for wing planform and scale?

LiftRoll calculates a value of average Cl as a percentage of maximum Cl ==> Can this percentage be applied in the spreadsheet to correct the idealized (2-dimensional) values of Cl provided by the Profili / XFoil analysis, in order to account for planform???

Concerning Reynolds number, what value of Re should be used in the airfoil analysis to obtain the baseline values of Cl and Cd? ==> OK to use an average value?

Any assistance much appreciated!

[Ollie]

PC Soar is a free down load that makes fewer simplifying assumptions and provides more accurate results. See:
http://my.athenet.net/~atkron95/pcsoar.htm

[Sail 'n Soar]

LiftRoll calculates a value of average Cl as a percentage of maximum Cl ==> Can this percentage be applied in the spreadsheet to correct the idealized (2-dimensional) values of Cl provided by the Profili / XFoil analysis, in order to account for planform???

Yes

Concerning Reynolds number, what value of Re should be used in the airfoil analysis to obtain the baseline values of Cl and Cd? ==> OK to use an average value?

Because the Cd decreases non-linearly with Re, if you are using some representative value, choose something between the average and tip Re.

My application is design of a thermal glider, where I want to get some idea of sink and lift / drag ratio (glide slope) in order to assess various combinations of wing planform, airfoil, aspect ratio, and wing loading.

If your tip cords aren't too small and your taper ratios not too severe, these approximations should be good enough for what you are trying to do.

In terms of L/D vs. wing loading, one interesting fact is that for similar planforms where just the scale changes, e.g., same AR, planform, etc., L/D vs alpha is a function of weight alone. (As the size goes down the wing loading and associated flying speed for L=W goes up such that Re is a constant.) Of course, for this case the sink will go as the square root of the weight.