NACA M20 AIRFOIL (m20-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M20 AIRFOIL (m20-il) Reynolds number: 500,000 Max Cl/Cd: 108.66 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m20-il-500000.txt Download as CSV file: xf-m20-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3595 0.10245 0.09965 0.0043 0.7335 0.0177
-7.750 -0.3524 0.09933 0.09651 0.0015 0.7262 0.0184
-7.500 -0.3474 0.09642 0.09357 -0.0025 0.7193 0.0186
-7.250 -0.3381 0.09314 0.09027 -0.0080 0.7128 0.0187
-7.000 -0.2534 0.07507 0.07223 -0.0153 0.6887 0.0190
-6.750 -0.2464 0.07157 0.06872 -0.0156 0.6829 0.0193
-6.500 -0.2384 0.06816 0.06529 -0.0170 0.6770 0.0196
-6.250 -0.2291 0.06484 0.06191 -0.0188 0.6719 0.0202
-6.000 -0.2180 0.06127 0.05833 -0.0211 0.6663 0.0209
-5.750 -0.2052 0.05751 0.05451 -0.0238 0.6610 0.0219
-5.500 -0.1823 0.05374 0.05062 -0.0298 0.6565 0.0234
-5.250 -0.1594 0.04966 0.04643 -0.0344 0.6512 0.0236
-5.000 -0.1390 0.04572 0.04238 -0.0368 0.6459 0.0236
-4.750 -0.1281 0.03997 0.03655 -0.0380 0.6416 0.0241
-4.500 -0.1138 0.03695 0.03353 -0.0379 0.6357 0.0245
-4.250 -0.0955 0.03421 0.03071 -0.0387 0.6302 0.0251
-4.000 -0.0923 0.04828 0.04445 -0.0404 0.6363 0.0254
-3.750 -0.0670 0.04581 0.04186 -0.0418 0.6306 0.0265
-3.500 -0.0386 0.04331 0.03919 -0.0434 0.6252 0.0284
-3.250 0.0028 0.04158 0.03721 -0.0453 0.6192 0.0301
-3.000 0.0312 0.03898 0.03440 -0.0460 0.6139 0.0302
-2.750 0.0491 0.03432 0.02963 -0.0468 0.6089 0.0313
-2.500 0.0733 0.03262 0.02786 -0.0472 0.6030 0.0321
-2.250 0.0991 0.03102 0.02611 -0.0475 0.5978 0.0331
-2.000 0.1267 0.02931 0.02428 -0.0479 0.5920 0.0347
-1.750 0.1619 0.02854 0.02324 -0.0478 0.5862 0.0388
-1.500 0.1891 0.02487 0.01920 -0.0478 0.5817 0.0398
-1.250 0.2140 0.02316 0.01745 -0.0481 0.5759 0.0410
-1.000 0.2403 0.02216 0.01635 -0.0483 0.5705 0.0424
-0.750 0.2679 0.02113 0.01518 -0.0484 0.5653 0.0447
-0.500 0.2997 0.02112 0.01495 -0.0480 0.5594 0.0500
-0.250 0.3258 0.01839 0.01195 -0.0482 0.5545 0.0526
0.000 0.3530 0.01770 0.01123 -0.0484 0.5492 0.0548
0.250 0.3812 0.01703 0.01046 -0.0485 0.5438 0.0585
0.500 0.4117 0.01783 0.01099 -0.0481 0.5388 0.0634
0.750 0.4378 0.01541 0.00853 -0.0485 0.5337 0.0683
1.000 0.4678 0.01381 0.00663 -0.0480 0.5287 0.0565
1.250 0.4952 0.01383 0.00661 -0.0483 0.5238 0.0695
1.500 0.5252 0.01175 0.00417 -0.0474 0.5195 0.0438
1.750 0.5531 0.01137 0.00376 -0.0473 0.5144 0.0438
2.000 0.5805 0.01108 0.00339 -0.0472 0.5098 0.0435
2.250 0.6081 0.01085 0.00316 -0.0472 0.5049 0.0439
2.500 0.6359 0.01068 0.00301 -0.0472 0.5001 0.0453
2.750 0.6635 0.01061 0.00290 -0.0472 0.4958 0.0473
3.000 0.6914 0.01057 0.00285 -0.0473 0.4914 0.0494
3.250 0.7194 0.01050 0.00280 -0.0474 0.4866 0.0537
3.500 0.7472 0.01047 0.00284 -0.0475 0.4822 0.0771
3.750 0.8100 0.00885 0.00315 -0.0557 0.4766 1.0000
4.000 0.8364 0.00893 0.00321 -0.0555 0.4715 1.0000
4.250 0.8624 0.00905 0.00328 -0.0553 0.4666 1.0000
4.500 0.8886 0.00917 0.00338 -0.0551 0.4621 1.0000
4.750 0.9149 0.00926 0.00349 -0.0549 0.4574 1.0000
5.000 0.9410 0.00939 0.00359 -0.0547 0.4528 1.0000
5.250 0.9670 0.00949 0.00370 -0.0545 0.4454 1.0000
5.750 1.0189 0.00967 0.00386 -0.0541 0.4263 1.0000
6.000 1.0448 0.00980 0.00399 -0.0540 0.4172 1.0000
6.250 1.0705 0.00994 0.00411 -0.0538 0.4030 1.0000
6.500 1.0960 0.01011 0.00427 -0.0536 0.3893 1.0000
6.750 1.1214 0.01032 0.00445 -0.0534 0.3695 1.0000
7.000 1.1442 0.01088 0.00479 -0.0530 0.3201 1.0000
7.250 1.1458 0.01443 0.00699 -0.0511 0.0956 1.0000
7.500 1.1583 0.01615 0.00838 -0.0496 0.0292 1.0000
7.750 1.1781 0.01690 0.00922 -0.0487 0.0248 1.0000
8.000 1.1983 0.01754 0.00994 -0.0479 0.0231 1.0000
8.250 1.2168 0.01831 0.01077 -0.0469 0.0213 1.0000
8.500 1.2321 0.01931 0.01187 -0.0457 0.0198 1.0000
8.750 1.2403 0.02083 0.01351 -0.0437 0.0186 1.0000
9.000 1.2473 0.02223 0.01503 -0.0416 0.0181 1.0000
9.250 1.2527 0.02349 0.01637 -0.0395 0.0177 1.0000
9.500 1.2565 0.02518 0.01816 -0.0380 0.0173 1.0000
9.750 1.2608 0.02714 0.02022 -0.0371 0.0168 1.0000
10.000 1.2652 0.02925 0.02242 -0.0364 0.0162 1.0000
10.250 1.2696 0.03143 0.02467 -0.0359 0.0156 1.0000
10.500 1.2729 0.03373 0.02704 -0.0354 0.0151 1.0000
10.750 1.2743 0.03623 0.02960 -0.0347 0.0147 1.0000
11.000 1.2747 0.03880 0.03222 -0.0340 0.0144 1.0000
11.250 1.2740 0.04141 0.03488 -0.0331 0.0141 1.0000
11.500 1.2730 0.04389 0.03739 -0.0317 0.0138 1.0000
11.750 1.2751 0.04578 0.03928 -0.0292 0.0135 1.0000
12.000 1.2861 0.04687 0.04039 -0.0264 0.0132 1.0000
12.250 1.2952 0.04855 0.04219 -0.0256 0.0131 1.0000
12.500 1.3067 0.05000 0.04371 -0.0243 0.0129 1.0000
12.750 1.3173 0.05159 0.04541 -0.0231 0.0127 1.0000
13.000 1.3264 0.05338 0.04730 -0.0222 0.0123 1.0000
13.250 1.3355 0.05521 0.04925 -0.0212 0.0119 1.0000
13.500 1.3475 0.05694 0.05110 -0.0197 0.0118 1.0000
13.750 1.3602 0.05890 0.05322 -0.0180 0.0118 1.0000
14.000 1.3705 0.06146 0.05596 -0.0163 0.0120 1.0000
14.250 1.3783 0.06484 0.05954 -0.0147 0.0125 1.0000
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Polar data table (+)
Polar graphs
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