NASA SC(2)-0503 AIRFOIL (sc20503-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0503 AIRFOIL (sc20503-il) Reynolds number: 200,000 Max Cl/Cd: 52.34 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20503-il-200000.txt Download as CSV file: xf-sc20503-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0503 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.6775 0.12966 0.12604 0.0311 1.0000 0.0186
-10.250 -0.6733 0.12589 0.12227 0.0294 1.0000 0.0191
-10.000 -0.6692 0.12213 0.11853 0.0277 1.0000 0.0196
-9.750 -0.6649 0.11842 0.11484 0.0258 1.0000 0.0201
-9.500 -0.6606 0.11478 0.11122 0.0236 1.0000 0.0205
-9.250 -0.6563 0.11123 0.10769 0.0211 1.0000 0.0210
-9.000 -0.6520 0.10771 0.10420 0.0183 1.0000 0.0213
-8.750 -0.6475 0.10417 0.10069 0.0152 1.0000 0.0215
-8.500 -0.6429 0.10054 0.09709 0.0121 1.0000 0.0217
-8.250 -0.6379 0.09671 0.09329 0.0083 1.0000 0.0218
-8.000 -0.6276 0.09219 0.08878 0.0026 1.0000 0.0220
-7.750 -0.6148 0.08719 0.08376 -0.0040 1.0000 0.0220
-7.500 -0.6009 0.08228 0.07880 -0.0094 1.0000 0.0221
-7.250 -0.5860 0.07748 0.07384 -0.0140 1.0000 0.0222
-7.000 -0.5701 0.07277 0.06901 -0.0180 1.0000 0.0222
-6.750 -0.5647 0.06452 0.06075 -0.0217 1.0000 0.0231
-6.500 -0.5552 0.05981 0.05605 -0.0226 1.0000 0.0243
-6.250 -0.5386 0.05558 0.05174 -0.0249 1.0000 0.0257
-6.000 -0.5172 0.05119 0.04719 -0.0278 1.0000 0.0275
-5.750 -0.4923 0.04677 0.04254 -0.0307 1.0000 0.0298
-5.500 -0.4597 0.04331 0.03867 -0.0332 1.0000 0.0337
-5.250 -0.4298 0.03855 0.03336 -0.0359 1.0000 0.0359
-5.000 -0.4085 0.03424 0.02909 -0.0374 1.0000 0.0392
-3.750 -0.2516 0.01876 0.01133 -0.0400 1.0000 0.0271
-3.500 -0.2211 0.01682 0.00902 -0.0391 1.0000 0.0205
-3.250 -0.1919 0.01474 0.00668 -0.0388 1.0000 0.0198
-3.000 -0.1633 0.01311 0.00493 -0.0387 1.0000 0.0242
-2.750 -0.1346 0.01195 0.00379 -0.0389 1.0000 0.0325
-2.500 -0.1036 0.01070 0.00247 -0.0395 1.0000 0.0485
-2.250 -0.0807 0.00746 0.00207 -0.0388 1.0000 0.8073
-2.000 -0.0788 0.00712 0.00197 -0.0314 1.0000 0.9102
-1.750 -0.0642 0.00673 0.00159 -0.0278 1.0000 1.0000
-1.500 -0.0334 0.00672 0.00142 -0.0288 1.0000 1.0000
-1.250 -0.0032 0.00672 0.00128 -0.0296 1.0000 1.0000
-1.000 0.0265 0.00673 0.00121 -0.0303 1.0000 1.0000
-0.750 0.0556 0.00674 0.00118 -0.0308 1.0000 1.0000
-0.500 0.0844 0.00677 0.00119 -0.0313 1.0000 1.0000
-0.250 0.1129 0.00680 0.00124 -0.0316 1.0000 1.0000
0.000 0.1411 0.00684 0.00133 -0.0318 1.0000 1.0000
0.250 0.1692 0.00689 0.00146 -0.0320 1.0000 1.0000
0.500 0.1971 0.00695 0.00166 -0.0321 1.0000 1.0000
0.750 0.2249 0.00701 0.00190 -0.0322 1.0000 1.0000
1.000 0.2844 0.01105 0.00280 -0.0385 0.0441 1.0000
1.250 0.3115 0.01235 0.00414 -0.0381 0.0319 1.0000
1.500 0.3390 0.01363 0.00545 -0.0376 0.0278 1.0000
1.750 0.3663 0.01525 0.00712 -0.0373 0.0194 1.0000
2.000 0.3950 0.01735 0.00946 -0.0366 0.0198 1.0000
2.250 0.4109 0.00785 0.00069 -0.0325 0.0263 1.0000
2.500 0.4640 0.02328 0.01654 -0.0327 0.0796 1.0000
2.750 0.4901 0.02556 0.01903 -0.0325 0.0641 1.0000
3.000 0.5168 0.02811 0.02213 -0.0321 0.0531 1.0000
3.250 0.5388 0.03256 0.02650 -0.0326 0.0478 1.0000
3.500 0.5680 0.03388 0.02856 -0.0316 0.0388 1.0000
3.750 0.5884 0.03774 0.03245 -0.0319 0.0354 1.0000
4.000 0.6076 0.04386 0.03890 -0.0322 0.0337 1.0000
4.250 0.6409 0.04508 0.04072 -0.0315 0.0294 1.0000
4.500 0.6641 0.04905 0.04495 -0.0320 0.0270 1.0000
4.750 0.6847 0.05314 0.04924 -0.0328 0.0251 1.0000
5.000 0.7016 0.05729 0.05353 -0.0336 0.0236 1.0000
5.250 0.7125 0.06225 0.05851 -0.0341 0.0223 1.0000
5.500 0.7171 0.07103 0.06736 -0.0360 0.0214 1.0000
5.750 0.7332 0.07557 0.07208 -0.0376 0.0213 1.0000
6.000 0.7485 0.08018 0.07692 -0.0398 0.0213 1.0000
6.250 0.7629 0.08482 0.08168 -0.0426 0.0211 1.0000
6.500 0.7764 0.08949 0.08644 -0.0461 0.0210 1.0000
6.750 0.7885 0.09427 0.09130 -0.0507 0.0208 1.0000
7.000 0.7987 0.09907 0.09614 -0.0561 0.0206 1.0000
7.250 0.8045 0.10369 0.10076 -0.0615 0.0204 1.0000
7.500 0.8065 0.10798 0.10503 -0.0658 0.0202 1.0000
7.750 0.8087 0.11214 0.10916 -0.0692 0.0199 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0503 AIRFOIL (sc20503-il)