durabol
Well-Known Member
I don't know if I accomplished anything with my examination of some airfoil with the computer program X-foil other than stir up more questions (and getting a head-ache) but I'll post my results for review. I got fairly similar results for reflexed (high CLmax) and unreflexed (low drag count) airfoils. Is it better to counter the pitching of the wing by reflexing the wing or with the tail or doesn't it matter? Even the "turbulent flow" 23012 doesn't have a much higher drag count and it has low pitching. The "drag clean/flaps" numbers are corrected for CLmax with: 5lbs drag at cruise=.1 change in CLmax=10 drag points. See math below:
-area=800lbs/(66.15^2*2*.5*.00237)=77.1sq.ft
-CL at cruise=800/(.5*.00237*220.5^2*77.1)=.18; CL at 700=.16; CL at 650=.15
CD=drag / (.5 *.00237 * 220.5^2 * 77.1sq.ft)
-40:1 lift/drag= 1000lbs with 25lbs drag, CD=.0056
-20:1, 1000lbs with 50lbs drag, CD=.011
-10:1, 1000lbs with 100lb drag, CD=.022
-CD=.02, very clean design with glider like fuselage
-CD=.02=fuselage.004 + tail.003 + engine.008 + wing.005 (wing is 25% of parasite drag), not certain how accurate these numbers are
-drag=.02 * .5 * .00237 * fps^2 * wing area
-drag 100mph=39lbs + some induced/vortex drag
-drag 120mph=56lbs + some induced/vortex drag
-drag 150mph=89lbs + little induced/vortex drag
-determine how much more area at 45mph with CL .1 less, calculate drag at 150mph for new wing area, subtract totals
-area of .1 CL lower=area=800lbs/(66.15^2*1.9*.5*.00237)=81.2sq.ft
-drag at 150mph with bigger wing area=.02 * .5 * .00237 * 220.5^2 * 81.2sq.ft=93.6lbs, 93.6-88.7=4.9lbs extra drag for CL lower by .1
-drag at 150mph with 5 more drag points: CDwing=.0055; .0205 * .5 * .00237 * 220.5^2 * 77.1sq.ft=91.1lbs, 91.1-88.7=2.4lbs
-In general at 150mph cruise/800lb gross airplane: 5lbs drag at cruise=.1 change in CLmax=10 change in drag points, this means a high CLmax is more important than getting the absolute lowest drag count but some low CL airfoils have large improvement with flaps so have to consider flap performance.
Brock
-area=800lbs/(66.15^2*2*.5*.00237)=77.1sq.ft
-CL at cruise=800/(.5*.00237*220.5^2*77.1)=.18; CL at 700=.16; CL at 650=.15
CD=drag / (.5 *.00237 * 220.5^2 * 77.1sq.ft)
-40:1 lift/drag= 1000lbs with 25lbs drag, CD=.0056
-20:1, 1000lbs with 50lbs drag, CD=.011
-10:1, 1000lbs with 100lb drag, CD=.022
-CD=.02, very clean design with glider like fuselage
-CD=.02=fuselage.004 + tail.003 + engine.008 + wing.005 (wing is 25% of parasite drag), not certain how accurate these numbers are
-drag=.02 * .5 * .00237 * fps^2 * wing area
-drag 100mph=39lbs + some induced/vortex drag
-drag 120mph=56lbs + some induced/vortex drag
-drag 150mph=89lbs + little induced/vortex drag
-determine how much more area at 45mph with CL .1 less, calculate drag at 150mph for new wing area, subtract totals
-area of .1 CL lower=area=800lbs/(66.15^2*1.9*.5*.00237)=81.2sq.ft
-drag at 150mph with bigger wing area=.02 * .5 * .00237 * 220.5^2 * 81.2sq.ft=93.6lbs, 93.6-88.7=4.9lbs extra drag for CL lower by .1
-drag at 150mph with 5 more drag points: CDwing=.0055; .0205 * .5 * .00237 * 220.5^2 * 77.1sq.ft=91.1lbs, 91.1-88.7=2.4lbs
-In general at 150mph cruise/800lb gross airplane: 5lbs drag at cruise=.1 change in CLmax=10 change in drag points, this means a high CLmax is more important than getting the absolute lowest drag count but some low CL airfoils have large improvement with flaps so have to consider flap performance.
Brock
