NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M16 AIRFOIL (m16-il) Reynolds number: 200,000 Max Cl/Cd: 70.26 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m16-il-200000.txt Download as CSV file: xf-m16-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-6.750 -0.4324 0.08696 0.08394 -0.0074 1.0000 0.0276
-6.500 -0.4277 0.08115 0.07818 -0.0060 1.0000 0.0283
-6.250 -0.4156 0.07704 0.07409 -0.0065 1.0000 0.0291
-6.000 -0.3992 0.07321 0.07026 -0.0089 1.0000 0.0301
-5.750 -0.3791 0.06941 0.06645 -0.0123 1.0000 0.0316
-5.500 -0.3539 0.06560 0.06258 -0.0169 1.0000 0.0344
-5.250 -0.3021 0.06294 0.05959 -0.0264 0.9385 0.0382
-5.000 -0.2769 0.06048 0.05686 -0.0279 0.9068 0.0385
-4.750 -0.2639 0.05480 0.05100 -0.0284 0.8859 0.0392
-4.500 -0.2553 0.05076 0.04693 -0.0273 0.8672 0.0405
-4.250 -0.2392 0.04807 0.04413 -0.0269 0.8508 0.0424
-4.000 -0.2171 0.04555 0.04143 -0.0273 0.8366 0.0463
-3.750 -0.1726 0.04543 0.04062 -0.0288 0.8238 0.0520
-3.500 -0.1612 0.03938 0.03461 -0.0292 0.8130 0.0543
-3.250 -0.1399 0.03717 0.03228 -0.0291 0.8019 0.0576
-3.000 -0.1041 0.03539 0.02990 -0.0297 0.7912 0.0666
-2.750 -0.0856 0.03253 0.02713 -0.0298 0.7810 0.0704
-2.500 -0.0544 0.03108 0.02514 -0.0298 0.7724 0.0811
-2.250 -0.0315 0.02869 0.02278 -0.0301 0.7623 0.0847
-2.000 -0.0026 0.02715 0.02088 -0.0301 0.7535 0.0963
-1.750 0.0245 0.02601 0.01944 -0.0300 0.7453 0.1095
-1.500 0.0513 0.02464 0.01795 -0.0302 0.7360 0.1237
-1.250 0.0780 0.02338 0.01651 -0.0302 0.7281 0.1379
-1.000 0.1046 0.02202 0.01506 -0.0303 0.7194 0.1532
-0.750 0.1442 0.01928 0.01127 -0.0287 0.7127 0.0806
-0.500 0.1724 0.01755 0.00942 -0.0285 0.7047 0.0741
-0.250 0.2011 0.01670 0.00838 -0.0283 0.6966 0.0743
0.000 0.2296 0.01572 0.00712 -0.0278 0.6893 0.0693
0.250 0.2582 0.01517 0.00644 -0.0277 0.6809 0.0672
0.500 0.2855 0.01469 0.00584 -0.0273 0.6739 0.0667
0.750 0.3133 0.01410 0.00528 -0.0272 0.6652 0.0674
1.000 0.3403 0.01367 0.00485 -0.0269 0.6581 0.0697
1.250 0.3680 0.01338 0.00460 -0.0269 0.6496 0.0736
1.500 0.3955 0.01321 0.00441 -0.0267 0.6420 0.0788
1.750 0.4231 0.01303 0.00426 -0.0266 0.6340 0.0931
2.000 0.4607 0.01080 0.00412 -0.0285 0.6257 1.0000
2.250 0.4872 0.01093 0.00407 -0.0281 0.6182 1.0000
2.500 0.5144 0.01105 0.00416 -0.0280 0.6095 1.0000
2.750 0.5408 0.01121 0.00419 -0.0277 0.6025 1.0000
3.000 0.5682 0.01133 0.00433 -0.0277 0.5934 1.0000
3.250 0.5952 0.01148 0.00444 -0.0275 0.5857 1.0000
3.500 0.6223 0.01162 0.00459 -0.0274 0.5772 1.0000
3.750 0.6495 0.01177 0.00477 -0.0273 0.5688 1.0000
4.000 0.6765 0.01193 0.00489 -0.0271 0.5612 1.0000
4.250 0.7040 0.01209 0.00513 -0.0271 0.5519 1.0000
4.500 0.7311 0.01222 0.00533 -0.0270 0.5429 1.0000
4.750 0.7576 0.01207 0.00518 -0.0265 0.5237 1.0000
5.000 0.7840 0.01196 0.00505 -0.0261 0.5029 1.0000
5.250 0.8108 0.01193 0.00508 -0.0258 0.4800 1.0000
5.500 0.8368 0.01191 0.00498 -0.0254 0.4303 1.0000
5.750 0.8529 0.01490 0.00613 -0.0257 0.0929 1.0000
6.000 0.8756 0.01652 0.00752 -0.0255 0.0425 1.0000
6.250 0.8994 0.01758 0.00871 -0.0253 0.0356 1.0000
6.500 0.9204 0.01910 0.01041 -0.0248 0.0328 1.0000
6.750 0.9416 0.02037 0.01179 -0.0242 0.0317 1.0000
7.000 0.9615 0.02182 0.01335 -0.0232 0.0309 1.0000
7.250 0.9813 0.02335 0.01494 -0.0221 0.0300 1.0000
7.500 1.0019 0.02475 0.01638 -0.0212 0.0278 1.0000
7.750 1.0230 0.02647 0.01816 -0.0201 0.0270 1.0000
8.000 1.0459 0.02856 0.02038 -0.0190 0.0274 1.0000
8.250 1.0703 0.03123 0.02326 -0.0178 0.0290 1.0000
8.500 1.0939 0.03529 0.02755 -0.0167 0.0319 1.0000
14.750 0.9391 0.17785 0.17431 -0.0617 0.0422 1.0000
15.000 0.9402 0.18353 0.17997 -0.0652 0.0399 1.0000
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Polar data table (+)
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