S1210 12% (s1210-il)
S1210 12% - Selig S1210 high lift low Reynolds number airfoil
| Details | Dat file | Parser | |
|
(s1210-il) S1210 12% Selig S1210 high lift low Reynolds number airfoil Max thickness 12% at 21.4% chord. Max camber 6.7% at 51.1% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Selig format |
S1210 12%
1.00000 0.00000
0.99837 0.00101
0.99398 0.00397
0.98753 0.00832
0.97908 0.01317
0.96811 0.01811
0.95437 0.02328
0.93796 0.02874
0.91898 0.03443
0.89754 0.04032
0.87376 0.04637
0.84779 0.05254
0.81980 0.05879
0.78997 0.06506
0.75851 0.07130
0.72561 0.07747
0.69151 0.08349
0.65642 0.08932
0.62058 0.09490
0.58423 0.10016
0.54763 0.10505
0.51105 0.10948
0.47473 0.11335
0.43891 0.11653
0.40378 0.11892
0.36955 0.12046
0.33652 0.12091
0.30456 0.12000
0.27347 0.11784
0.24341 0.11462
0.21445 0.11047
0.18681 0.10556
0.16069 0.09994
0.13622 0.09362
0.11351 0.08672
0.09269 0.07932
0.07388 0.07149
0.05719 0.06332
0.04282 0.05484
0.03068 0.04593
0.02054 0.03672
0.01239 0.02755
0.00626 0.01866
0.00217 0.01030
0.00016 0.00277
0.00023 -0.00345
0.00337 -0.00773
0.01034 -0.01070
0.02071 -0.01324
0.03417 -0.01529
0.05052 -0.01685
0.06959 -0.01786
0.09118 -0.01830
0.11512 -0.01810
0.14119 -0.01715
0.16911 -0.01524
0.19906 -0.01183
0.23157 -0.00697
0.26670 -0.00124
0.30427 0.00504
0.34404 0.01158
0.38575 0.01814
0.42909 0.02446
0.47370 0.03032
0.51919 0.03551
0.56515 0.03986
0.61113 0.04320
0.65666 0.04543
0.70127 0.04646
0.74446 0.04625
0.78575 0.04479
0.82465 0.04214
0.86071 0.03837
0.89349 0.03364
0.92255 0.02809
0.94754 0.02192
0.96791 0.01530
0.98299 0.00890
0.99284 0.00390
0.99828 0.00095
1.00000 0.00000
|
No parser warnings |
Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
|
Polars for S1210 12% (s1210-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| s1210-il | 50,000 | 9 | 14.9 at α=2.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1210-il | 50,000 | 5 | 42.1 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1210-il | 100,000 | 9 | 59.3 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1210-il | 100,000 | 5 | 66 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1210-il | 200,000 | 9 | 86.3 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1210-il | 200,000 | 5 | 88 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1210-il | 500,000 | 9 | 121.4 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1210-il | 500,000 | 5 | 115 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1210-il | 1,000,000 | 9 | 148 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1210-il | 1,000,000 | 5 | 135.9 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||