NACA M21 AIRFOIL (m21-il) Xfoil prediction polar at RE=500,000 Ncrit=5
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Airfoil: NACA M21 AIRFOIL (m21-il) Reynolds number: 500,000 Max Cl/Cd: 105.82 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m21-il-500000-n5.txt Download as CSV file: xf-m21-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M21 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3130 0.09657 0.09312 -0.0178 0.6466 0.0200
-8.750 -0.3102 0.09278 0.08933 -0.0200 0.6419 0.0205
-8.500 -0.3259 0.08489 0.08145 -0.0265 0.6388 0.0217
-8.250 -0.3199 0.08242 0.07894 -0.0284 0.6336 0.0219
-8.000 -0.3106 0.08034 0.07684 -0.0297 0.6287 0.0222
-7.750 -0.2997 0.07779 0.07427 -0.0318 0.6235 0.0224
-7.500 -0.2890 0.07479 0.07121 -0.0343 0.6188 0.0227
-7.250 -0.2775 0.07153 0.06789 -0.0369 0.6147 0.0231
-7.000 -0.2649 0.06801 0.06433 -0.0396 0.6104 0.0234
-6.750 -0.2513 0.06424 0.06048 -0.0422 0.6058 0.0239
-5.250 -0.1846 0.01961 0.01329 -0.0496 0.5860 0.0307
-5.000 -0.1611 0.01756 0.01085 -0.0492 0.5818 0.0313
-4.750 -0.1343 0.01680 0.00995 -0.0491 0.5770 0.0317
-4.500 -0.1070 0.01624 0.00927 -0.0490 0.5721 0.0321
-4.250 -0.0796 0.01569 0.00858 -0.0489 0.5678 0.0326
-4.000 -0.0520 0.01515 0.00791 -0.0489 0.5635 0.0331
-3.750 -0.0242 0.01461 0.00725 -0.0489 0.5586 0.0337
-3.500 0.0037 0.01412 0.00662 -0.0489 0.5540 0.0343
-3.250 0.0316 0.01368 0.00605 -0.0489 0.5500 0.0350
-3.000 0.0598 0.01328 0.00555 -0.0489 0.5456 0.0356
-2.750 0.0879 0.01296 0.00513 -0.0489 0.5409 0.0362
-2.500 0.1158 0.01266 0.00478 -0.0490 0.5365 0.0373
-2.250 0.1441 0.01249 0.00459 -0.0490 0.5325 0.0382
-2.000 0.1724 0.01231 0.00438 -0.0491 0.5281 0.0392
-1.750 0.2005 0.01212 0.00415 -0.0491 0.5235 0.0403
-1.500 0.2284 0.01195 0.00391 -0.0491 0.5194 0.0415
-1.250 0.2566 0.01178 0.00369 -0.0491 0.5155 0.0426
-1.000 0.2845 0.01161 0.00354 -0.0492 0.5110 0.0443
-0.750 0.3127 0.01155 0.00346 -0.0492 0.5066 0.0460
-0.250 0.3690 0.01142 0.00327 -0.0494 0.4987 0.0502
0.000 0.3970 0.01134 0.00322 -0.0495 0.4943 0.0528
0.250 0.4251 0.01133 0.00319 -0.0495 0.4899 0.0551
0.500 0.4531 0.01131 0.00314 -0.0496 0.4861 0.0576
0.750 0.4812 0.01125 0.00307 -0.0497 0.4823 0.0596
1.000 0.5091 0.01121 0.00306 -0.0498 0.4779 0.0624
1.250 0.5371 0.01123 0.00307 -0.0499 0.4735 0.0653
1.500 0.5651 0.01125 0.00307 -0.0500 0.4696 0.0681
1.750 0.5931 0.01122 0.00305 -0.0501 0.4657 0.0704
2.000 0.6206 0.01117 0.00303 -0.0501 0.4615 0.0733
2.250 0.6483 0.01119 0.00305 -0.0502 0.4573 0.0762
2.500 0.6760 0.01121 0.00306 -0.0503 0.4533 0.0791
2.750 0.7039 0.01122 0.00308 -0.0504 0.4491 0.0812
3.000 0.7314 0.01122 0.00309 -0.0505 0.4450 0.0840
3.250 0.7587 0.01125 0.00313 -0.0506 0.4410 0.0879
3.500 0.7863 0.01129 0.00319 -0.0507 0.4370 0.0915
3.750 0.8140 0.01133 0.00325 -0.0508 0.4325 0.0947
4.000 0.8413 0.01137 0.00331 -0.0509 0.4282 0.0993
4.250 0.8684 0.01145 0.00340 -0.0509 0.4243 0.1061
4.500 0.8959 0.01149 0.00350 -0.0511 0.4206 0.1178
4.750 0.9147 0.01057 0.00365 -0.0499 0.4166 0.6589
6.000 1.1371 0.01101 0.00486 -0.0693 0.3925 1.0000
6.250 1.1617 0.01113 0.00501 -0.0689 0.3884 1.0000
6.500 1.1858 0.01129 0.00518 -0.0684 0.3844 1.0000
6.750 1.2095 0.01148 0.00538 -0.0679 0.3801 1.0000
7.000 1.2329 0.01167 0.00558 -0.0673 0.3749 1.0000
7.250 1.2561 0.01187 0.00578 -0.0667 0.3677 1.0000
7.500 1.2783 0.01212 0.00601 -0.0660 0.3602 1.0000
7.750 1.3009 0.01233 0.00625 -0.0653 0.3527 1.0000
8.000 1.3222 0.01262 0.00652 -0.0645 0.3458 1.0000
8.250 1.3442 0.01286 0.00678 -0.0637 0.3379 1.0000
8.500 1.3641 0.01319 0.00710 -0.0626 0.3280 1.0000
8.750 1.3835 0.01355 0.00745 -0.0615 0.3177 1.0000
9.000 1.4027 0.01389 0.00779 -0.0603 0.3074 1.0000
9.250 1.4196 0.01434 0.00821 -0.0588 0.2945 1.0000
9.500 1.4332 0.01493 0.00873 -0.0568 0.2765 1.0000
9.750 1.4440 0.01561 0.00934 -0.0544 0.2582 1.0000
10.000 1.4488 0.01648 0.01011 -0.0512 0.2346 1.0000
10.250 1.4390 0.01772 0.01120 -0.0456 0.2069 1.0000
10.500 1.4256 0.01955 0.01287 -0.0407 0.1766 1.0000
10.750 1.4110 0.02216 0.01530 -0.0371 0.1431 1.0000
11.000 1.3841 0.02629 0.01917 -0.0338 0.0959 1.0000
11.250 1.3577 0.03090 0.02358 -0.0314 0.0565 1.0000
11.500 1.3395 0.03503 0.02761 -0.0299 0.0309 1.0000
11.750 1.3354 0.03794 0.03053 -0.0290 0.0233 1.0000
12.000 1.3356 0.04047 0.03310 -0.0284 0.0204 1.0000
12.250 1.3353 0.04307 0.03576 -0.0278 0.0185 1.0000
12.500 1.3368 0.04554 0.03830 -0.0273 0.0174 1.0000
12.750 1.3381 0.04808 0.04091 -0.0269 0.0165 1.0000
13.000 1.3385 0.05079 0.04370 -0.0265 0.0158 1.0000
13.250 1.3381 0.05367 0.04665 -0.0263 0.0152 1.0000
13.500 1.3362 0.05677 0.04982 -0.0261 0.0145 1.0000
13.750 1.3352 0.05981 0.05294 -0.0260 0.0140 1.0000
14.000 1.3358 0.06270 0.05592 -0.0260 0.0136 1.0000
14.250 1.3356 0.06575 0.05905 -0.0260 0.0132 1.0000
14.500 1.3345 0.06894 0.06232 -0.0261 0.0128 1.0000
14.750 1.3332 0.07218 0.06563 -0.0263 0.0124 1.0000
15.000 1.3313 0.07553 0.06906 -0.0265 0.0120 1.0000
15.250 1.3280 0.07913 0.07274 -0.0268 0.0117 1.0000
15.500 1.3240 0.08284 0.07652 -0.0271 0.0114 1.0000
15.750 1.3187 0.08676 0.08052 -0.0276 0.0112 1.0000
16.000 1.3140 0.09069 0.08452 -0.0281 0.0110 1.0000
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Polar data table (+)
Polar graphs
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