NASA SC(2)-0410 AIRFOIL (sc20410-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0410 AIRFOIL (sc20410-il) Reynolds number: 50,000 Max Cl/Cd: 22.06 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20410-il-50000-n5.txt Download as CSV file: xf-sc20410-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0410 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.4978 0.13178 0.12402 -0.0119 1.0000 0.0682
-12.250 -0.5000 0.12787 0.12012 -0.0132 1.0000 0.0677
-11.750 -0.6629 0.12137 0.11327 -0.0101 1.0000 0.0606
-11.500 -0.6622 0.11639 0.10831 -0.0120 1.0000 0.0603
-11.250 -0.6642 0.11083 0.10278 -0.0144 1.0000 0.0601
-11.000 -0.6699 0.10427 0.09625 -0.0180 1.0000 0.0599
-10.750 -0.6797 0.09662 0.08861 -0.0228 1.0000 0.0597
-10.500 -0.6947 0.08910 0.08109 -0.0277 1.0000 0.0593
-10.250 -0.7138 0.08223 0.07417 -0.0321 1.0000 0.0590
-10.000 -0.7348 0.07630 0.06818 -0.0352 1.0000 0.0587
-9.750 -0.7559 0.07131 0.06307 -0.0368 1.0000 0.0585
-9.500 -0.7708 0.06642 0.05796 -0.0384 1.0000 0.0587
-9.250 -0.7805 0.06175 0.05295 -0.0392 1.0000 0.0592
-9.000 -0.7848 0.05724 0.04798 -0.0396 1.0000 0.0600
-8.750 -0.7805 0.05326 0.04362 -0.0397 1.0000 0.0611
-8.500 -0.7680 0.05032 0.04053 -0.0395 1.0000 0.0625
-8.250 -0.7540 0.04738 0.03735 -0.0393 1.0000 0.0639
-8.000 -0.7383 0.04459 0.03427 -0.0391 1.0000 0.0658
-7.750 -0.7209 0.04195 0.03120 -0.0388 1.0000 0.0691
-7.500 -0.7016 0.03925 0.02788 -0.0386 1.0000 0.0725
-7.250 -0.6810 0.03707 0.02569 -0.0380 1.0000 0.0754
-7.000 -0.6595 0.03537 0.02386 -0.0375 1.0000 0.0800
-6.750 -0.6366 0.03360 0.02168 -0.0368 1.0000 0.0852
-6.500 -0.6146 0.03198 0.02008 -0.0359 1.0000 0.0898
-6.250 -0.5920 0.03070 0.01871 -0.0349 1.0000 0.0965
-6.000 -0.5693 0.02944 0.01734 -0.0334 1.0000 0.1024
-5.750 -0.5476 0.02841 0.01632 -0.0321 1.0000 0.1104
-5.500 -0.5259 0.02741 0.01528 -0.0306 1.0000 0.1186
-5.250 -0.5047 0.02647 0.01432 -0.0293 1.0000 0.1289
-5.000 -0.4841 0.02548 0.01342 -0.0283 1.0000 0.1420
-4.750 -0.4631 0.02445 0.01247 -0.0275 1.0000 0.1592
-4.500 -0.4419 0.02323 0.01153 -0.0271 1.0000 0.1879
-4.250 -0.4204 0.02165 0.01052 -0.0274 1.0000 0.2539
-4.000 -0.4043 0.02002 0.01030 -0.0259 1.0000 0.4368
-3.500 -0.3747 0.02134 0.01204 -0.0168 1.0000 0.6544
-3.250 -0.3568 0.02194 0.01252 -0.0136 1.0000 0.6926
-3.000 -0.3420 0.02255 0.01306 -0.0094 1.0000 0.7247
-2.750 -0.3306 0.02316 0.01361 -0.0041 1.0000 0.7564
-2.500 -0.3213 0.02356 0.01398 0.0016 1.0000 0.7838
-2.250 -0.3096 0.02366 0.01403 0.0062 1.0000 0.8067
-2.000 -0.2947 0.02353 0.01382 0.0095 1.0000 0.8246
-1.750 -0.2781 0.02329 0.01350 0.0120 1.0000 0.8397
-1.500 -0.2597 0.02300 0.01312 0.0138 1.0000 0.8523
-1.250 -0.2397 0.02272 0.01276 0.0149 1.0000 0.8633
-1.000 -0.2191 0.02241 0.01238 0.0159 1.0000 0.8722
-0.750 -0.1967 0.02214 0.01205 0.0163 1.0000 0.8792
-0.500 -0.1738 0.02189 0.01175 0.0165 1.0000 0.8852
-0.250 -0.1497 0.02172 0.01155 0.0163 1.0000 0.8914
0.000 -0.1267 0.02150 0.01132 0.0165 1.0000 0.8971
0.250 -0.1021 0.02139 0.01119 0.0162 1.0000 0.9025
0.500 -0.0773 0.02129 0.01111 0.0158 1.0000 0.9074
0.750 -0.0526 0.02122 0.01106 0.0155 1.0000 0.9120
1.000 -0.0271 0.02122 0.01110 0.0148 1.0000 0.9167
1.250 -0.0017 0.02124 0.01119 0.0142 1.0000 0.9213
1.500 0.0232 0.02129 0.01132 0.0136 1.0000 0.9261
1.750 0.0577 0.02147 0.01160 0.0112 0.9948 0.9304
2.000 0.1068 0.02172 0.01199 0.0062 0.9793 0.9331
2.250 0.1547 0.02186 0.01230 0.0017 0.9623 0.9356
2.500 0.2065 0.02187 0.01252 -0.0031 0.9414 0.9375
2.750 0.2651 0.02144 0.01234 -0.0078 0.9046 0.9384
3.000 0.3194 0.02066 0.01181 -0.0107 0.8594 0.9397
3.250 0.3606 0.01992 0.01125 -0.0110 0.8020 0.9424
3.500 0.3963 0.01936 0.01070 -0.0101 0.7060 0.9457
3.750 0.4285 0.01966 0.00999 -0.0082 0.4698 0.9483
4.000 0.4439 0.02115 0.01038 -0.0062 0.2946 0.9530
4.250 0.4657 0.02250 0.01114 -0.0061 0.2117 0.9581
4.500 0.4918 0.02360 0.01200 -0.0064 0.1724 0.9634
4.750 0.5201 0.02462 0.01292 -0.0070 0.1495 0.9695
5.000 0.5510 0.02563 0.01395 -0.0079 0.1329 0.9759
5.250 0.5827 0.02674 0.01507 -0.0089 0.1206 0.9838
5.500 0.6134 0.02781 0.01623 -0.0097 0.1101 1.0000
5.750 0.6440 0.02920 0.01776 -0.0103 0.1019 1.0000
6.000 0.6729 0.03060 0.01916 -0.0111 0.0944 1.0000
6.250 0.7031 0.03237 0.02123 -0.0118 0.0882 1.0000
6.500 0.7312 0.03398 0.02296 -0.0125 0.0821 1.0000
6.750 0.7590 0.03619 0.02541 -0.0131 0.0779 1.0000
7.000 0.7850 0.03845 0.02816 -0.0135 0.0729 1.0000
7.250 0.8100 0.04068 0.03061 -0.0139 0.0694 1.0000
7.500 0.8338 0.04333 0.03335 -0.0145 0.0670 1.0000
7.750 0.8511 0.04687 0.03768 -0.0142 0.0644 1.0000
8.000 0.8661 0.05047 0.04183 -0.0140 0.0618 1.0000
8.250 0.8784 0.05431 0.04616 -0.0139 0.0603 1.0000
8.500 0.8866 0.05856 0.05084 -0.0137 0.0596 1.0000
8.750 0.8899 0.06315 0.05583 -0.0137 0.0591 1.0000
9.000 0.8880 0.06804 0.06107 -0.0139 0.0588 1.0000
9.250 0.8802 0.07320 0.06654 -0.0144 0.0586 1.0000
9.500 0.8634 0.07886 0.07245 -0.0154 0.0588 1.0000
9.750 0.8372 0.08594 0.07971 -0.0186 0.0596 1.0000
10.000 0.8118 0.09523 0.08910 -0.0259 0.0607 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0410 AIRFOIL (sc20410-il)