NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M16 AIRFOIL (m16-il) Reynolds number: 50,000 Max Cl/Cd: 38.13 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m16-il-50000-n5.txt Download as CSV file: xf-m16-il-50000-n5.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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4683 0.11194 0.10584 0.0119 1.0000 0.0617
-8.000 -0.4639 0.10910 0.10306 0.0091 1.0000 0.0649
-7.750 -0.4621 0.10724 0.10129 0.0041 1.0000 0.0669
-7.500 -0.4547 0.10559 0.09966 -0.0032 1.0000 0.0678
-7.250 -0.4452 0.09937 0.09353 -0.0019 1.0000 0.0694
-7.000 -0.4347 0.09470 0.08883 -0.0012 1.0000 0.0727
-6.750 -0.4234 0.09107 0.08522 -0.0039 1.0000 0.0762
-6.500 -0.4088 0.08826 0.08239 -0.0098 1.0000 0.0807
-6.000 -0.3807 0.07994 0.07410 -0.0149 1.0000 0.0853
-5.750 -0.3631 0.07670 0.07081 -0.0181 1.0000 0.0941
-5.500 -0.3462 0.07264 0.06675 -0.0209 1.0000 0.0994
-5.250 -0.3190 0.07075 0.06462 -0.0270 1.0000 0.1104
-4.750 -0.2782 0.06363 0.05736 -0.0316 1.0000 0.1250
-4.500 -0.2686 0.05873 0.05261 -0.0303 1.0000 0.1297
-4.250 -0.2497 0.05576 0.04958 -0.0320 0.9896 0.1461
-3.750 -0.1804 0.04841 0.04180 -0.0406 0.9469 0.1898
-3.500 -0.1187 0.04330 0.03573 -0.0471 0.9313 0.0925
-3.250 -0.0840 0.04028 0.03229 -0.0493 0.9149 0.0913
-3.000 -0.0465 0.03737 0.02864 -0.0509 0.8998 0.0813
-2.750 -0.0177 0.03497 0.02587 -0.0514 0.8849 0.0794
-2.500 0.0111 0.03295 0.02343 -0.0516 0.8707 0.0780
-2.250 0.0396 0.03119 0.02123 -0.0515 0.8573 0.0773
-2.000 0.0679 0.02979 0.01935 -0.0512 0.8444 0.0797
-1.750 0.0943 0.02864 0.01788 -0.0510 0.8319 0.0857
-1.500 0.1221 0.02747 0.01636 -0.0506 0.8201 0.0863
-1.250 0.1502 0.02641 0.01490 -0.0501 0.8091 0.0855
-1.000 0.1778 0.02550 0.01365 -0.0494 0.7987 0.0852
-0.750 0.2069 0.02476 0.01258 -0.0492 0.7874 0.0853
-0.500 0.2363 0.02413 0.01169 -0.0490 0.7767 0.0858
-0.250 0.2640 0.02359 0.01096 -0.0484 0.7669 0.0868
0.000 0.2904 0.02315 0.01036 -0.0477 0.7572 0.0884
0.250 0.3164 0.02282 0.00988 -0.0472 0.7468 0.0908
0.500 0.3417 0.02256 0.00948 -0.0465 0.7374 0.0945
0.750 0.3667 0.02232 0.00917 -0.0457 0.7283 0.1000
1.000 0.3925 0.02217 0.00901 -0.0454 0.7182 0.1116
1.250 0.4177 0.02187 0.00890 -0.0448 0.7092 0.1534
1.500 0.4580 0.02005 0.00873 -0.0470 0.7001 1.0000
1.750 0.4836 0.02037 0.00886 -0.0467 0.6900 1.0000
2.000 0.5086 0.02066 0.00902 -0.0460 0.6812 1.0000
2.250 0.5337 0.02097 0.00923 -0.0455 0.6721 1.0000
2.500 0.5591 0.02134 0.00956 -0.0452 0.6623 1.0000
2.750 0.5842 0.02164 0.00983 -0.0445 0.6541 1.0000
3.000 0.6095 0.02203 0.01023 -0.0442 0.6444 1.0000
3.250 0.6349 0.02245 0.01069 -0.0440 0.6349 1.0000
3.500 0.6598 0.02275 0.01099 -0.0432 0.6273 1.0000
3.750 0.6852 0.02326 0.01166 -0.0431 0.6169 1.0000
4.000 0.7104 0.02371 0.01222 -0.0428 0.6078 1.0000
4.250 0.7355 0.02408 0.01269 -0.0422 0.5995 1.0000
4.500 0.7605 0.02466 0.01347 -0.0421 0.5893 1.0000
4.750 0.7855 0.02514 0.01416 -0.0416 0.5805 1.0000
5.000 0.8105 0.02560 0.01483 -0.0411 0.5715 1.0000
5.250 0.8350 0.02626 0.01577 -0.0410 0.5612 1.0000
5.500 0.8599 0.02677 0.01655 -0.0404 0.5525 1.0000
5.750 0.8847 0.02719 0.01733 -0.0397 0.5425 1.0000
6.000 0.9056 0.02577 0.01604 -0.0362 0.5036 1.0000
6.250 0.9254 0.02451 0.01481 -0.0328 0.4458 1.0000
6.500 0.9392 0.02463 0.01427 -0.0298 0.2918 1.0000
6.750 0.9351 0.02931 0.01696 -0.0286 0.0698 1.0000
7.000 0.9466 0.03173 0.01939 -0.0276 0.0556 1.0000
7.250 0.9574 0.03392 0.02179 -0.0266 0.0481 1.0000
7.500 0.9655 0.03620 0.02422 -0.0254 0.0432 1.0000
7.750 0.9711 0.03857 0.02675 -0.0240 0.0411 1.0000
8.000 0.9785 0.04063 0.02903 -0.0224 0.0395 1.0000
8.250 0.9864 0.04263 0.03122 -0.0207 0.0377 1.0000
8.500 0.9950 0.04454 0.03326 -0.0188 0.0354 1.0000
8.750 1.0115 0.04653 0.03518 -0.0165 0.0326 1.0000
9.000 1.0439 0.04816 0.03714 -0.0147 0.0312 1.0000
9.250 1.0753 0.05065 0.04005 -0.0133 0.0304 1.0000
9.500 1.0981 0.05373 0.04349 -0.0121 0.0300 1.0000
9.750 1.1119 0.05701 0.04717 -0.0110 0.0293 1.0000
10.000 1.1189 0.06038 0.05095 -0.0097 0.0286 1.0000
10.250 1.1205 0.06386 0.05479 -0.0085 0.0279 1.0000
10.500 1.1163 0.06732 0.05857 -0.0072 0.0275 1.0000
10.750 1.1079 0.07093 0.06247 -0.0063 0.0273 1.0000
11.000 1.0970 0.07491 0.06672 -0.0062 0.0273 1.0000
11.250 1.0841 0.07933 0.07139 -0.0069 0.0273 1.0000
11.500 1.0695 0.08420 0.07650 -0.0084 0.0275 1.0000
11.750 1.0537 0.08959 0.08209 -0.0106 0.0277 1.0000
12.000 1.0368 0.09552 0.08821 -0.0135 0.0280 1.0000
12.250 1.0195 0.10200 0.09485 -0.0172 0.0283 1.0000
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Polar data table (+)
Polar graphs
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