S1221 w/o flap (s1221-il)
S1221 w/o flap - Selig S1221 airfoil
Details | Dat file | Parser | |
(s1221-il) S1221 w/o flap Selig S1221 airfoil Max thickness 12.1% at 21.8% chord. Max camber 4.9% at 49.1% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Selig format |
S1221 w/o flap 1.00182 0.01052 1.00010 0.01125 0.99539 0.01347 0.98834 0.01676 0.97911 0.02033 0.96725 0.02380 0.95255 0.02730 0.94744 0.02839 0.87368 0.04200 0.79952 0.05285 0.72514 0.06224 0.65074 0.07145 0.64181 0.07259 0.60457 0.07717 0.56677 0.08157 0.52865 0.08577 0.49051 0.08975 0.45264 0.09349 0.41537 0.09693 0.37902 0.09992 0.34385 0.10232 0.31005 0.10396 0.27776 0.10469 0.24715 0.10447 0.21836 0.10298 0.19125 0.10011 0.16569 0.09601 0.14166 0.09077 0.11918 0.08461 0.09838 0.07771 0.07937 0.07022 0.06223 0.06222 0.04705 0.05386 0.03388 0.04528 0.02281 0.03663 0.01391 0.02808 0.00733 0.01969 0.00297 0.01152 0.00065 0.00384 0.00037 -0.00287 0.00271 -0.00774 0.00859 -0.01117 0.01809 -0.01422 0.03068 -0.01687 0.04619 -0.01910 0.06442 -0.02087 0.08523 -0.02216 0.10841 -0.02290 0.13379 -0.02304 0.16116 -0.02248 0.19026 -0.02101 0.22124 -0.01814 0.25456 -0.01393 0.29026 -0.00894 0.32811 -0.00346 0.36788 0.00227 0.40929 0.00802 0.45202 0.01360 0.49574 0.01880 0.54006 0.02346 0.58460 0.02742 0.62895 0.03058 0.65127 0.03183 0.72628 0.03528 0.80124 0.03663 0.87610 0.03380 0.95081 0.02472 0.96969 0.02084 0.98452 0.01668 0.99445 0.01326 1.00003 0.01119 1.00181 0.01052 |
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Similar airfoils
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Polars for S1221 w/o flap (s1221-il)
Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
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s1221-il | 50,000 | 9 | 27.1 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
s1221-il | 50,000 | 5 | 36.6 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
s1221-il | 100,000 | 9 | 45.2 at α=3.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
s1221-il | 100,000 | 5 | 54.5 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
s1221-il | 200,000 | 9 | 70 at α=11.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
s1221-il | 200,000 | 5 | 76.5 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
s1221-il | 500,000 | 9 | 107.5 at α=9.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
s1221-il | 500,000 | 5 | 110 at α=8.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
s1221-il | 1,000,000 | 9 | 137.9 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
s1221-il | 1,000,000 | 5 | 135.1 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
Reynolds number calculator |