NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: NACA M16 AIRFOIL (m16-il) Reynolds number: 100,000 Max Cl/Cd: 53.56 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m16-il-100000.txt Download as CSV file: xf-m16-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M16 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4820 0.10608 0.10174 0.0163 1.0000 0.0464
-7.750 -0.4765 0.10282 0.09853 0.0138 1.0000 0.0481
-7.500 -0.4715 0.10000 0.09577 0.0098 1.0000 0.0499
-7.250 -0.4591 0.09833 0.09409 0.0006 1.0000 0.0513
-7.000 -0.4398 0.09693 0.09259 -0.0089 1.0000 0.0519
-6.750 -0.4364 0.08898 0.08478 -0.0055 1.0000 0.0532
-6.500 -0.4271 0.08420 0.08005 -0.0039 1.0000 0.0551
-6.250 -0.4131 0.08038 0.07623 -0.0059 1.0000 0.0576
-6.000 -0.3955 0.07668 0.07252 -0.0095 1.0000 0.0607
-5.750 -0.3577 0.07534 0.07090 -0.0204 1.0000 0.0654
-5.500 -0.3411 0.06985 0.06540 -0.0228 1.0000 0.0667
-5.250 -0.3306 0.06495 0.06063 -0.0218 1.0000 0.0692
-5.000 -0.3084 0.06161 0.05725 -0.0244 1.0000 0.0756
-4.750 -0.2783 0.05828 0.05374 -0.0300 1.0000 0.0816
-4.500 -0.2766 0.05584 0.05140 -0.0286 1.0000 0.0843
-4.250 -0.2243 0.05228 0.04744 -0.0373 0.9857 0.0951
-4.000 -0.1891 0.04776 0.04289 -0.0413 0.9718 0.1036
-3.750 -0.1493 0.04424 0.03913 -0.0460 0.9568 0.1169
-3.500 -0.1100 0.04222 0.03667 -0.0498 0.9402 0.1368
-3.250 -0.0840 0.03968 0.03396 -0.0506 0.9229 0.1510
-3.000 -0.0621 0.03720 0.03136 -0.0504 0.9069 0.1659
-2.750 -0.0417 0.03498 0.02905 -0.0497 0.8920 0.1827
-2.500 -0.0216 0.03317 0.02711 -0.0488 0.8782 0.2121
-2.250 -0.0022 0.03140 0.02522 -0.0479 0.8654 0.2516
-2.000 0.0172 0.03006 0.02374 -0.0466 0.8535 0.3034
-1.750 0.0352 0.02764 0.02134 -0.0449 0.8429 0.3427
-1.500 0.0575 0.02567 0.01930 -0.0440 0.8314 0.3804
-1.250 0.0855 0.02427 0.01768 -0.0439 0.8205 0.3974
-1.000 0.1546 0.02419 0.01577 -0.0458 0.8112 0.1326
-0.750 0.1828 0.02282 0.01400 -0.0446 0.8018 0.1162
-0.500 0.2124 0.02196 0.01271 -0.0440 0.7912 0.1068
-0.250 0.2393 0.02099 0.01162 -0.0434 0.7814 0.1043
0.000 0.2651 0.02026 0.01071 -0.0422 0.7729 0.1029
0.250 0.2929 0.01966 0.01003 -0.0420 0.7623 0.1032
0.500 0.3198 0.01920 0.00949 -0.0414 0.7525 0.1054
0.750 0.3445 0.01870 0.00897 -0.0402 0.7446 0.1107
1.000 0.3715 0.01845 0.00875 -0.0400 0.7339 0.1234
1.250 0.3979 0.01822 0.00851 -0.0394 0.7245 0.1363
1.500 0.4229 0.01780 0.00823 -0.0384 0.7166 0.1861
1.750 0.4657 0.01608 0.00802 -0.0410 0.7059 1.0000
2.000 0.4913 0.01637 0.00817 -0.0405 0.6965 1.0000
2.250 0.5159 0.01658 0.00824 -0.0395 0.6884 1.0000
2.500 0.5423 0.01691 0.00856 -0.0395 0.6779 1.0000
2.750 0.5680 0.01722 0.00883 -0.0391 0.6687 1.0000
3.000 0.5931 0.01746 0.00900 -0.0383 0.6605 1.0000
3.250 0.6195 0.01785 0.00942 -0.0383 0.6501 1.0000
3.500 0.6453 0.01818 0.00979 -0.0379 0.6412 1.0000
3.750 0.6709 0.01846 0.01007 -0.0373 0.6327 1.0000
4.000 0.6972 0.01889 0.01060 -0.0373 0.6224 1.0000
4.250 0.7230 0.01924 0.01099 -0.0368 0.6140 1.0000
4.500 0.7488 0.01957 0.01144 -0.0364 0.6048 1.0000
4.750 0.7751 0.02006 0.01209 -0.0364 0.5947 1.0000
5.000 0.8008 0.02038 0.01251 -0.0359 0.5861 1.0000
5.250 0.8250 0.02003 0.01222 -0.0342 0.5680 1.0000
5.500 0.8476 0.01873 0.01084 -0.0312 0.5394 1.0000
5.750 0.8714 0.01777 0.00992 -0.0290 0.5086 1.0000
6.000 0.8945 0.01670 0.00882 -0.0267 0.4460 1.0000
6.250 0.9031 0.01978 0.00972 -0.0255 0.0970 1.0000
6.500 0.9219 0.02191 0.01165 -0.0248 0.0670 1.0000
6.750 0.9421 0.02340 0.01333 -0.0240 0.0599 1.0000
7.000 0.9585 0.02520 0.01527 -0.0229 0.0547 1.0000
7.250 0.9761 0.02668 0.01686 -0.0217 0.0507 1.0000
7.500 0.9937 0.02836 0.01853 -0.0201 0.0490 1.0000
7.750 1.0144 0.03022 0.02035 -0.0186 0.0479 1.0000
8.000 1.0391 0.03235 0.02250 -0.0174 0.0476 1.0000
8.250 1.0653 0.03478 0.02503 -0.0164 0.0474 1.0000
8.500 1.0898 0.03807 0.02837 -0.0158 0.0457 1.0000
8.750 1.1128 0.04102 0.03166 -0.0148 0.0457 1.0000
9.000 1.1330 0.04374 0.03514 -0.0129 0.0496 1.0000
9.250 1.1509 0.04835 0.04020 -0.0116 0.0539 1.0000
9.500 1.1657 0.05257 0.04521 -0.0094 0.0633 1.0000
9.750 1.1741 0.05946 0.05316 -0.0070 0.0882 1.0000
10.750 1.0149 0.07074 0.06603 0.0042 0.1099 1.0000
11.000 0.9878 0.07619 0.07158 0.0034 0.1093 1.0000
11.250 0.9616 0.08224 0.07770 0.0018 0.1087 1.0000
11.500 0.9345 0.08897 0.08448 -0.0004 0.1082 1.0000
11.750 0.9068 0.09624 0.09174 -0.0032 0.1079 1.0000
12.000 0.8768 0.10422 0.09974 -0.0070 0.1076 1.0000
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Polar data table (+)
Polar graphs
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