NACA M17 AIRFOIL (m17-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M17 AIRFOIL (m17-il) Reynolds number: 100,000 Max Cl/Cd: 55.17 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m17-il-100000.txt Download as CSV file: xf-m17-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M17 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 -0.4342 0.09935 0.09499 -0.0008 1.0000 0.0662
-7.750 -0.4320 0.09642 0.09213 -0.0050 1.0000 0.0686
-7.500 -0.4280 0.09498 0.09068 -0.0154 1.0000 0.0704
-7.250 -0.4190 0.09121 0.08689 -0.0212 1.0000 0.0713
-7.000 -0.4090 0.08487 0.08068 -0.0149 1.0000 0.0739
-6.750 -0.3969 0.08126 0.07710 -0.0160 1.0000 0.0769
-6.500 -0.3833 0.07779 0.07364 -0.0195 1.0000 0.0807
-6.250 -0.3588 0.07795 0.07342 -0.0315 1.0000 0.0853
-6.000 -0.3539 0.07127 0.06699 -0.0297 1.0000 0.0866
-5.750 -0.3613 0.06922 0.06505 -0.0266 0.9981 0.0877
-5.500 -0.3263 0.06460 0.06036 -0.0316 0.9864 0.0942
-5.250 -0.2846 0.05990 0.05540 -0.0402 0.9729 0.1027
-5.000 -0.2494 0.05613 0.05150 -0.0446 0.9593 0.1111
-4.750 -0.2165 0.05224 0.04739 -0.0491 0.9444 0.1181
-4.500 -0.1842 0.05025 0.04498 -0.0523 0.9286 0.1304
-4.250 -0.1671 0.04672 0.04154 -0.0515 0.9138 0.1362
-4.000 -0.1462 0.04450 0.03911 -0.0516 0.8989 0.1485
-3.750 -0.1254 0.04260 0.03697 -0.0514 0.8850 0.1617
-3.500 -0.1053 0.04075 0.03493 -0.0508 0.8720 0.1760
-3.250 -0.0858 0.03877 0.03281 -0.0500 0.8599 0.1912
-3.000 -0.0666 0.03680 0.03074 -0.0489 0.8490 0.2079
-2.750 -0.0479 0.03515 0.02895 -0.0477 0.8389 0.2358
-2.500 -0.0292 0.03359 0.02734 -0.0465 0.8277 0.2740
-2.250 -0.0116 0.03194 0.02560 -0.0451 0.8179 0.3195
-2.000 0.0060 0.02960 0.02325 -0.0433 0.8095 0.3536
-1.750 0.0781 0.02768 0.01947 -0.0465 0.8003 0.1268
-1.500 0.1063 0.02577 0.01711 -0.0453 0.7927 0.1097
-1.250 0.1343 0.02445 0.01548 -0.0448 0.7833 0.1055
-1.000 0.1614 0.02357 0.01437 -0.0443 0.7749 0.1065
-0.750 0.1890 0.02262 0.01313 -0.0435 0.7669 0.1045
-0.500 0.2171 0.02188 0.01222 -0.0432 0.7584 0.1040
-0.250 0.2438 0.02121 0.01145 -0.0424 0.7512 0.1054
0.000 0.2715 0.02073 0.01095 -0.0423 0.7425 0.1086
0.250 0.2976 0.02027 0.01039 -0.0414 0.7360 0.1136
0.500 0.3243 0.01991 0.01017 -0.0413 0.7273 0.1264
0.750 0.3505 0.01949 0.00975 -0.0404 0.7211 0.1466
1.000 0.4346 0.01708 0.00928 -0.0515 0.7123 1.0000
1.250 0.4586 0.01729 0.00928 -0.0504 0.7061 1.0000
1.500 0.4842 0.01767 0.00958 -0.0504 0.6976 1.0000
1.750 0.5083 0.01789 0.00967 -0.0494 0.6915 1.0000
2.000 0.5337 0.01833 0.01008 -0.0494 0.6834 1.0000
2.250 0.5579 0.01857 0.01022 -0.0485 0.6773 1.0000
2.500 0.5828 0.01904 0.01070 -0.0484 0.6694 1.0000
2.750 0.6072 0.01933 0.01094 -0.0477 0.6632 1.0000
3.000 0.6317 0.01981 0.01144 -0.0474 0.6558 1.0000
3.250 0.6561 0.02015 0.01176 -0.0468 0.6492 1.0000
3.500 0.6804 0.02061 0.01225 -0.0463 0.6423 1.0000
3.750 0.7046 0.02101 0.01269 -0.0458 0.6350 1.0000
4.000 0.7287 0.02142 0.01312 -0.0452 0.6285 1.0000
4.250 0.7526 0.02186 0.01362 -0.0446 0.6204 1.0000
4.500 0.7763 0.02220 0.01401 -0.0438 0.6125 1.0000
4.750 0.8005 0.02222 0.01407 -0.0426 0.6037 1.0000
5.000 0.8237 0.02254 0.01447 -0.0418 0.5930 1.0000
5.250 0.8485 0.02248 0.01442 -0.0405 0.5857 1.0000
5.500 0.8719 0.02290 0.01498 -0.0399 0.5758 1.0000
5.750 0.8956 0.02327 0.01549 -0.0392 0.5668 1.0000
6.000 0.9208 0.02328 0.01556 -0.0381 0.5589 1.0000
6.250 0.9438 0.02373 0.01620 -0.0375 0.5482 1.0000
6.500 0.9683 0.02384 0.01644 -0.0365 0.5388 1.0000
6.750 0.9931 0.02295 0.01563 -0.0345 0.5209 1.0000
7.000 1.0181 0.02122 0.01380 -0.0318 0.4937 1.0000
7.250 1.0417 0.02025 0.01285 -0.0298 0.4665 1.0000
7.500 1.0634 0.01960 0.01232 -0.0280 0.4295 1.0000
7.750 1.0781 0.01954 0.01196 -0.0254 0.3210 1.0000
8.000 1.0677 0.02341 0.01417 -0.0219 0.1139 1.0000
8.250 1.0695 0.02604 0.01638 -0.0193 0.0797 1.0000
8.500 1.0765 0.02786 0.01824 -0.0171 0.0700 1.0000
8.750 1.0810 0.02973 0.02015 -0.0147 0.0642 1.0000
9.000 1.0871 0.03139 0.02187 -0.0124 0.0596 1.0000
9.250 1.0906 0.03319 0.02365 -0.0098 0.0570 1.0000
9.500 1.1030 0.03517 0.02555 -0.0075 0.0547 1.0000
9.750 1.1265 0.03680 0.02728 -0.0063 0.0529 1.0000
10.000 1.1502 0.03864 0.02922 -0.0054 0.0500 1.0000
10.250 1.1736 0.04075 0.03141 -0.0047 0.0474 1.0000
10.500 1.2019 0.04351 0.03432 -0.0044 0.0466 1.0000
10.750 1.2249 0.04657 0.03765 -0.0036 0.0465 1.0000
11.000 1.2410 0.04985 0.04133 -0.0023 0.0471 1.0000
11.250 1.2447 0.05330 0.04529 -0.0002 0.0480 1.0000
11.500 1.2431 0.05683 0.04924 0.0020 0.0487 1.0000
11.750 1.2354 0.06021 0.05297 0.0045 0.0491 1.0000
12.000 1.2233 0.06370 0.05677 0.0066 0.0496 1.0000
12.250 1.2083 0.06751 0.06086 0.0079 0.0501 1.0000
12.500 1.1911 0.07175 0.06537 0.0084 0.0507 1.0000
12.750 1.1722 0.07644 0.07030 0.0081 0.0513 1.0000
13.000 1.1522 0.08160 0.07567 0.0071 0.0519 1.0000
13.250 1.1316 0.08723 0.08149 0.0054 0.0526 1.0000
13.500 1.1111 0.09334 0.08776 0.0032 0.0533 1.0000
13.750 1.0928 0.09985 0.09439 0.0008 0.0541 1.0000
14.000 1.0929 0.10554 0.10015 0.0002 0.0556 1.0000
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Polar data table (+)
Polar graphs
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