NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M13 AIRFOIL (m13-il) Reynolds number: 50,000 Max Cl/Cd: 40.57 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m13-il-50000-n5.txt Download as CSV file: xf-m13-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M13 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4599 0.12111 0.11441 0.0016 1.0000 0.0701
-8.750 -0.4601 0.11944 0.11282 -0.0018 1.0000 0.0712
-8.500 -0.4612 0.11780 0.11127 -0.0056 1.0000 0.0717
-8.250 -0.4468 0.11111 0.10459 -0.0038 1.0000 0.0735
-8.000 -0.4383 0.10708 0.10059 -0.0041 1.0000 0.0759
-7.750 -0.4330 0.10383 0.09735 -0.0056 1.0000 0.0784
-7.500 -0.4294 0.10095 0.09453 -0.0081 1.0000 0.0807
-7.250 -0.4236 0.09849 0.09212 -0.0136 1.0000 0.0828
-7.000 -0.4141 0.09630 0.08993 -0.0209 1.0000 0.0837
-6.750 -0.4042 0.09109 0.08479 -0.0207 1.0000 0.0848
-6.500 -0.3943 0.08649 0.08023 -0.0186 1.0000 0.0885
-6.250 -0.3819 0.08317 0.07690 -0.0219 1.0000 0.0936
-5.750 -0.3547 0.07568 0.06941 -0.0279 1.0000 0.1001
-5.500 -0.3403 0.07202 0.06574 -0.0292 1.0000 0.1035
-5.250 -0.3146 0.06985 0.06329 -0.0366 1.0000 0.1109
-5.000 -0.3028 0.06514 0.05867 -0.0361 1.0000 0.1130
-4.750 -0.2810 0.06297 0.05629 -0.0395 1.0000 0.1260
-4.500 -0.2738 0.05886 0.05233 -0.0369 1.0000 0.1364
-4.250 -0.2595 0.05584 0.04928 -0.0373 1.0000 0.1486
-3.750 -0.1957 0.04687 0.03927 -0.0432 1.0000 0.0713
-3.500 -0.1843 0.04442 0.03684 -0.0423 1.0000 0.0749
-3.250 -0.1691 0.04238 0.03462 -0.0417 1.0000 0.0783
-3.000 -0.1332 0.03900 0.03076 -0.0447 0.9931 0.0750
-2.750 -0.0866 0.03563 0.02658 -0.0487 0.9823 0.0714
-2.500 -0.0450 0.03285 0.02327 -0.0520 0.9713 0.0711
-2.250 -0.0052 0.03084 0.02103 -0.0553 0.9601 0.0768
-2.000 0.0380 0.02888 0.01855 -0.0585 0.9499 0.0794
-1.750 0.0791 0.02707 0.01625 -0.0609 0.9386 0.0792
-1.500 0.1184 0.02560 0.01437 -0.0629 0.9264 0.0794
-1.250 0.1561 0.02438 0.01279 -0.0645 0.9140 0.0803
-1.000 0.1931 0.02339 0.01146 -0.0657 0.9015 0.0818
-0.750 0.2291 0.02258 0.01033 -0.0667 0.8888 0.0842
-0.500 0.2614 0.02184 0.00950 -0.0673 0.8756 0.0882
-0.250 0.2915 0.02130 0.00881 -0.0674 0.8622 0.0948
0.000 0.3198 0.02082 0.00828 -0.0672 0.8486 0.1059
0.250 0.3473 0.02044 0.00795 -0.0670 0.8352 0.1387
0.500 0.3734 0.01963 0.00774 -0.0669 0.8224 0.2551
0.750 0.4100 0.01800 0.00741 -0.0678 0.8104 1.0000
1.000 0.4358 0.01823 0.00737 -0.0672 0.7971 1.0000
1.250 0.4613 0.01847 0.00741 -0.0665 0.7838 1.0000
1.500 0.4866 0.01872 0.00751 -0.0658 0.7708 1.0000
1.750 0.5120 0.01898 0.00767 -0.0652 0.7582 1.0000
2.000 0.5373 0.01926 0.00785 -0.0646 0.7459 1.0000
2.250 0.5627 0.01953 0.00806 -0.0639 0.7340 1.0000
2.500 0.5882 0.01981 0.00831 -0.0633 0.7228 1.0000
2.750 0.6135 0.02012 0.00861 -0.0626 0.7109 1.0000
3.000 0.6389 0.02047 0.00899 -0.0622 0.6986 1.0000
3.250 0.6643 0.02083 0.00939 -0.0617 0.6867 1.0000
3.500 0.6898 0.02119 0.00984 -0.0612 0.6751 1.0000
3.750 0.7154 0.02153 0.01024 -0.0606 0.6642 1.0000
4.000 0.7408 0.02190 0.01071 -0.0600 0.6527 1.0000
4.250 0.7658 0.02233 0.01129 -0.0595 0.6403 1.0000
4.500 0.7907 0.02277 0.01193 -0.0590 0.6281 1.0000
4.750 0.8156 0.02321 0.01254 -0.0584 0.6160 1.0000
5.000 0.8405 0.02362 0.01316 -0.0577 0.6041 1.0000
5.250 0.8657 0.02401 0.01375 -0.0570 0.5924 1.0000
5.500 0.8905 0.02445 0.01450 -0.0562 0.5801 1.0000
5.750 0.9148 0.02497 0.01535 -0.0555 0.5670 1.0000
6.000 0.9392 0.02549 0.01624 -0.0548 0.5538 1.0000
6.250 0.9558 0.02410 0.01478 -0.0500 0.4939 1.0000
6.500 0.9667 0.02383 0.01421 -0.0458 0.3885 1.0000
6.750 0.9643 0.02679 0.01532 -0.0424 0.1289 1.0000
7.000 0.9689 0.03018 0.01802 -0.0406 0.0698 1.0000
7.250 0.9800 0.03236 0.02030 -0.0389 0.0585 1.0000
7.500 0.9894 0.03454 0.02260 -0.0370 0.0522 1.0000
7.750 0.9994 0.03650 0.02474 -0.0352 0.0463 1.0000
8.000 1.0060 0.03884 0.02710 -0.0332 0.0427 1.0000
8.250 1.0205 0.04066 0.02919 -0.0311 0.0406 1.0000
8.500 1.0402 0.04255 0.03130 -0.0293 0.0382 1.0000
8.750 1.0607 0.04465 0.03361 -0.0279 0.0349 1.0000
9.000 1.0853 0.04748 0.03653 -0.0270 0.0325 1.0000
9.250 1.1084 0.05047 0.03993 -0.0259 0.0316 1.0000
9.500 1.1256 0.05384 0.04374 -0.0246 0.0311 1.0000
9.750 1.1368 0.05743 0.04777 -0.0232 0.0308 1.0000
10.000 1.1416 0.06109 0.05186 -0.0216 0.0304 1.0000
10.250 1.1407 0.06477 0.05595 -0.0199 0.0300 1.0000
10.500 1.1342 0.06841 0.05994 -0.0182 0.0297 1.0000
10.750 1.1227 0.07203 0.06386 -0.0165 0.0296 1.0000
11.000 1.1089 0.07601 0.06811 -0.0158 0.0295 1.0000
11.250 1.0936 0.08047 0.07280 -0.0162 0.0295 1.0000
11.500 1.0770 0.08547 0.07802 -0.0176 0.0297 1.0000
11.750 1.0597 0.09104 0.08377 -0.0199 0.0299 1.0000
12.000 1.0421 0.09718 0.09008 -0.0230 0.0303 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M13 AIRFOIL (m13-il)