Ornithopter airfoil. (s1020-il)
Ornithopter airfoil. - Selig S1020 ornithopter airfoil
| Details | Dat file | Parser | |
|
(s1020-il) Ornithopter airfoil. Selig S1020 ornithopter airfoil Max thickness 15.1% at 33.8% chord. Max camber 4.6% at 57% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Selig format |
Ornithopter airfoil.
S1020
1.00000 0.00000
0.99637 0.00132
0.98633 0.00544
0.97138 0.01201
0.95221 0.01987
0.92858 0.02824
0.90037 0.03716
0.86802 0.04669
0.83206 0.05668
0.79303 0.06690
0.75150 0.07709
0.70800 0.08691
0.66299 0.09599
0.61686 0.10404
0.57001 0.11082
0.52281 0.11615
0.47565 0.11991
0.42893 0.12200
0.38308 0.12232
0.33841 0.12081
0.29521 0.11747
0.25379 0.11243
0.21447 0.10583
0.17759 0.09786
0.14348 0.08869
0.11246 0.07850
0.08479 0.06752
0.06069 0.05597
0.04036 0.04414
0.02399 0.03235
0.01172 0.02092
0.00368 0.01030
0.00006 0.00111
0.00210 -0.00597
0.01048 -0.01176
0.02453 -0.01716
0.04387 -0.02193
0.06822 -0.02596
0.09729 -0.02919
0.13077 -0.03158
0.16828 -0.03309
0.20943 -0.03370
0.25378 -0.03334
0.30086 -0.03199
0.35023 -0.02957
0.40159 -0.02597
0.45480 -0.02141
0.50939 -0.01639
0.56466 -0.01126
0.61990 -0.00630
0.67438 -0.00178
0.72731 0.00203
0.77785 0.00495
0.82515 0.00682
0.86836 0.00760
0.90663 0.00733
0.93918 0.00618
0.96531 0.00441
0.98443 0.00242
0.99609 0.00071
1.00000 0.00000
|
Dat file parser warnings
|
Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
|
Polars for Ornithopter airfoil. (s1020-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| s1020-il | 50,000 | 9 | 6 at α=8.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1020-il | 50,000 | 5 | 18.1 at α=8.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1020-il | 100,000 | 9 | 49.8 at α=10.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1020-il | 100,000 | 5 | 53.4 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1020-il | 200,000 | 9 | 82.4 at α=7.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1020-il | 200,000 | 5 | 81.5 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1020-il | 500,000 | 9 | 120 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1020-il | 500,000 | 5 | 111.2 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s1020-il | 1,000,000 | 9 | 146.8 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s1020-il | 1,000,000 | 5 | 127.2 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||