NACA M20 AIRFOIL (m20-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M20 AIRFOIL (m20-il) Reynolds number: 200,000 Max Cl/Cd: 78.64 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m20-il-200000.txt Download as CSV file: xf-m20-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M20 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.2543 0.10976 0.10674 -0.0101 0.8576 0.0271
-9.250 -0.2514 0.10683 0.10376 -0.0106 0.8463 0.0278
-9.000 -0.3503 0.11388 0.11090 -0.0013 0.9023 0.0260
-8.750 -0.3473 0.11130 0.10825 -0.0014 0.8800 0.0266
-8.500 -0.3429 0.10856 0.10544 -0.0021 0.8629 0.0272
-8.250 -0.3379 0.10578 0.10259 -0.0031 0.8488 0.0279
-8.000 -0.3326 0.10311 0.09986 -0.0046 0.8365 0.0288
-7.750 -0.3274 0.10132 0.09803 -0.0087 0.8240 0.0297
-7.500 -0.3219 0.10011 0.09677 -0.0158 0.8123 0.0300
-7.250 -0.3077 0.09783 0.09439 -0.0221 0.8028 0.0302
-7.000 -0.2980 0.09302 0.08952 -0.0240 0.7946 0.0304
-6.750 -0.2926 0.08757 0.08407 -0.0204 0.7865 0.0309
-6.500 -0.2828 0.08392 0.08037 -0.0201 0.7787 0.0315
-6.250 -0.2699 0.08064 0.07704 -0.0216 0.7702 0.0322
-6.000 -0.2556 0.07753 0.07384 -0.0237 0.7631 0.0332
-5.750 -0.2383 0.07434 0.07060 -0.0265 0.7549 0.0344
-5.500 -0.2191 0.07127 0.06741 -0.0294 0.7482 0.0362
-5.250 -0.1716 0.07067 0.06649 -0.0396 0.7399 0.0390
-5.000 -0.1522 0.06617 0.06184 -0.0419 0.7338 0.0394
-4.750 -0.1437 0.06133 0.05704 -0.0411 0.7269 0.0402
-4.500 -0.1284 0.05816 0.05381 -0.0410 0.7201 0.0412
-4.250 -0.1079 0.05542 0.05099 -0.0421 0.7134 0.0427
-4.000 -0.0839 0.05283 0.04827 -0.0436 0.7065 0.0451
-3.750 -0.0343 0.05302 0.04795 -0.0481 0.7003 0.0502
-3.500 -0.0157 0.04804 0.04291 -0.0492 0.6936 0.0511
-3.250 0.0004 0.04485 0.03967 -0.0490 0.6881 0.0523
-3.000 0.0223 0.04253 0.03731 -0.0494 0.6809 0.0543
-2.750 0.0477 0.04052 0.03513 -0.0500 0.6748 0.0573
-2.500 0.0896 0.03994 0.03401 -0.0514 0.6689 0.0638
-2.250 0.1079 0.03642 0.03054 -0.0517 0.6624 0.0654
-2.000 0.1305 0.03451 0.02851 -0.0517 0.6573 0.0677
-1.750 0.1579 0.03295 0.02684 -0.0522 0.6500 0.0722
-1.500 0.1913 0.03162 0.02505 -0.0524 0.6446 0.0787
-1.250 0.2145 0.02963 0.02305 -0.0526 0.6389 0.0812
-1.000 0.2418 0.02846 0.02176 -0.0528 0.6325 0.0877
-0.750 0.2698 0.02703 0.02005 -0.0528 0.6277 0.0951
-0.500 0.2972 0.02595 0.01889 -0.0530 0.6209 0.1016
-0.250 0.3255 0.02471 0.01742 -0.0530 0.6154 0.1098
0.000 0.3542 0.02403 0.01648 -0.0529 0.6105 0.1223
0.250 0.3808 0.02279 0.01526 -0.0531 0.6041 0.1289
0.500 0.4083 0.02189 0.01419 -0.0531 0.5992 0.1426
0.750 0.4358 0.02121 0.01340 -0.0532 0.5934 0.1669
1.000 0.4627 0.02037 0.01254 -0.0532 0.5877 0.1875
1.250 0.5026 0.01791 0.00912 -0.0515 0.5840 0.0729
1.500 0.5313 0.01714 0.00822 -0.0513 0.5787 0.0695
1.750 0.5595 0.01655 0.00753 -0.0512 0.5731 0.0683
2.000 0.5870 0.01613 0.00701 -0.0509 0.5688 0.0690
2.250 0.6143 0.01585 0.00678 -0.0508 0.5633 0.0710
2.500 0.6415 0.01564 0.00658 -0.0507 0.5582 0.0748
2.750 0.6683 0.01542 0.00636 -0.0504 0.5541 0.0823
3.000 0.6953 0.01535 0.00633 -0.0504 0.5488 0.0914
3.250 0.7187 0.01441 0.00626 -0.0498 0.5439 0.4894
3.500 0.7849 0.01366 0.00627 -0.0578 0.5390 1.0000
3.750 0.8109 0.01389 0.00650 -0.0576 0.5338 1.0000
4.000 0.8369 0.01407 0.00666 -0.0574 0.5288 1.0000
4.250 0.8629 0.01426 0.00676 -0.0571 0.5249 1.0000
4.500 0.8887 0.01453 0.00705 -0.0569 0.5206 1.0000
4.750 0.9145 0.01477 0.00733 -0.0567 0.5155 1.0000
5.000 0.9405 0.01496 0.00749 -0.0564 0.5114 1.0000
5.250 0.9663 0.01521 0.00774 -0.0562 0.5072 1.0000
5.500 0.9919 0.01548 0.00811 -0.0561 0.5016 1.0000
5.750 1.0178 0.01565 0.00828 -0.0558 0.4969 1.0000
6.000 1.0434 0.01583 0.00849 -0.0555 0.4910 1.0000
6.250 1.0689 0.01568 0.00833 -0.0550 0.4810 1.0000
6.500 1.0937 0.01555 0.00824 -0.0544 0.4688 1.0000
6.750 1.1185 0.01541 0.00812 -0.0539 0.4558 1.0000
7.000 1.1433 0.01539 0.00814 -0.0534 0.4439 1.0000
7.250 1.1676 0.01531 0.00814 -0.0529 0.4278 1.0000
7.500 1.1909 0.01524 0.00807 -0.0522 0.4021 1.0000
7.750 1.2126 0.01542 0.00820 -0.0513 0.3608 1.0000
8.000 1.2186 0.01743 0.00932 -0.0493 0.2148 1.0000
8.250 1.2053 0.02139 0.01225 -0.0458 0.0595 1.0000
8.500 1.2128 0.02301 0.01381 -0.0437 0.0423 1.0000
8.750 1.2234 0.02424 0.01518 -0.0420 0.0388 1.0000
9.000 1.2285 0.02571 0.01679 -0.0398 0.0365 1.0000
9.250 1.2269 0.02762 0.01883 -0.0374 0.0351 1.0000
9.500 1.2224 0.03024 0.02162 -0.0359 0.0337 1.0000
9.750 1.2220 0.03275 0.02426 -0.0351 0.0328 1.0000
10.000 1.2223 0.03532 0.02695 -0.0344 0.0318 1.0000
10.250 1.2208 0.03812 0.02986 -0.0337 0.0311 1.0000
10.500 1.2189 0.04096 0.03280 -0.0330 0.0305 1.0000
10.750 1.2178 0.04371 0.03564 -0.0322 0.0300 1.0000
11.000 1.2180 0.04625 0.03825 -0.0312 0.0295 1.0000
11.250 1.2203 0.04854 0.04059 -0.0300 0.0291 1.0000
11.500 1.2252 0.05046 0.04255 -0.0285 0.0287 1.0000
11.750 1.2341 0.05194 0.04405 -0.0266 0.0283 1.0000
12.000 1.2451 0.05328 0.04539 -0.0248 0.0274 1.0000
12.250 1.2647 0.05388 0.04591 -0.0219 0.0260 1.0000
12.500 1.2959 0.05425 0.04633 -0.0189 0.0258 1.0000
12.750 1.3194 0.05551 0.04774 -0.0172 0.0261 1.0000
13.000 1.3402 0.05721 0.04973 -0.0155 0.0273 1.0000
13.250 1.1678 0.05667 0.04966 -0.0058 0.0264 1.0000
13.500 1.1910 0.05835 0.05157 -0.0039 0.0272 1.0000
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Polar data table (+)
Polar graphs
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