NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA-M1 AIRFOIL (m1-il) Reynolds number: 50,000 Max Cl/Cd: 18.4 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m1-il-50000.txt Download as CSV file: xf-m1-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA-M1 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5470 0.09904 0.09251 0.0242 1.0000 0.3213
-8.250 -0.5506 0.09553 0.08905 0.0246 1.0000 0.3367
-8.000 -0.5589 0.09224 0.08582 0.0250 1.0000 0.3518
-7.750 -0.5374 0.08751 0.08107 0.0270 1.0000 0.3776
-7.500 -0.5255 0.08332 0.07689 0.0285 1.0000 0.4024
-7.000 -0.6488 0.08654 0.07991 0.0432 1.0000 0.4258
-6.750 -0.6477 0.08405 0.07747 0.0461 1.0000 0.4632
-6.250 -0.6138 0.07614 0.06953 0.0504 1.0000 0.5246
-6.000 -0.6477 0.05051 0.04223 -0.0088 1.0000 0.1545
-5.750 -0.6256 0.04508 0.03618 -0.0095 1.0000 0.1393
-5.500 -0.6035 0.04108 0.03168 -0.0094 1.0000 0.1379
-5.250 -0.5793 0.03730 0.02731 -0.0091 1.0000 0.1354
-5.000 -0.5527 0.03382 0.02306 -0.0086 1.0000 0.1319
-4.750 -0.5259 0.03103 0.01985 -0.0080 1.0000 0.1335
-4.500 -0.4988 0.02897 0.01728 -0.0073 1.0000 0.1418
-4.250 -0.4710 0.02661 0.01477 -0.0067 1.0000 0.1467
-4.000 -0.4423 0.02474 0.01269 -0.0059 1.0000 0.1540
-3.750 -0.4143 0.02308 0.01095 -0.0051 1.0000 0.1686
-3.500 -0.3865 0.02147 0.00947 -0.0045 1.0000 0.1919
-3.250 -0.3613 0.01963 0.00794 -0.0035 1.0000 0.2352
-3.000 -0.2520 0.01663 0.00742 -0.0102 1.0000 1.0000
-2.750 -0.2320 0.01620 0.00662 -0.0097 1.0000 1.0000
-2.500 -0.2115 0.01584 0.00600 -0.0091 1.0000 1.0000
-2.250 -0.1907 0.01556 0.00547 -0.0085 1.0000 1.0000
-2.000 -0.1698 0.01532 0.00502 -0.0077 1.0000 1.0000
-1.750 -0.1488 0.01513 0.00465 -0.0069 1.0000 1.0000
-1.500 -0.1276 0.01497 0.00431 -0.0060 1.0000 1.0000
-1.250 -0.1063 0.01485 0.00406 -0.0051 1.0000 1.0000
-1.000 -0.0850 0.01475 0.00386 -0.0041 1.0000 1.0000
-0.750 -0.0638 0.01467 0.00371 -0.0031 1.0000 1.0000
-0.500 -0.0426 0.01462 0.00359 -0.0021 1.0000 1.0000
-0.250 -0.0213 0.01459 0.00352 -0.0010 1.0000 1.0000
0.000 0.0000 0.01458 0.00350 0.0000 1.0000 1.0000
0.250 0.0213 0.01459 0.00352 0.0010 1.0000 1.0000
0.500 0.0425 0.01462 0.00359 0.0021 1.0000 1.0000
0.750 0.0638 0.01467 0.00371 0.0031 1.0000 1.0000
1.000 0.0850 0.01475 0.00386 0.0041 1.0000 1.0000
1.250 0.1063 0.01485 0.00406 0.0051 1.0000 1.0000
1.500 0.1276 0.01497 0.00431 0.0060 1.0000 1.0000
1.750 0.1488 0.01513 0.00465 0.0069 1.0000 1.0000
2.000 0.1698 0.01532 0.00502 0.0077 1.0000 1.0000
2.250 0.1908 0.01555 0.00547 0.0085 1.0000 1.0000
2.500 0.2115 0.01584 0.00599 0.0091 1.0000 1.0000
2.750 0.2320 0.01620 0.00662 0.0097 1.0000 1.0000
3.000 0.2521 0.01663 0.00742 0.0102 1.0000 1.0000
3.250 0.3613 0.01964 0.00794 0.0035 0.2352 1.0000
3.500 0.3865 0.02147 0.00947 0.0045 0.1919 1.0000
3.750 0.4143 0.02308 0.01095 0.0051 0.1686 1.0000
4.000 0.4423 0.02474 0.01269 0.0059 0.1539 1.0000
4.250 0.4710 0.02661 0.01477 0.0067 0.1467 1.0000
4.500 0.4988 0.02898 0.01728 0.0073 0.1418 1.0000
4.750 0.5259 0.03103 0.01985 0.0080 0.1336 1.0000
5.000 0.5527 0.03382 0.02306 0.0086 0.1319 1.0000
5.250 0.5793 0.03730 0.02731 0.0091 0.1354 1.0000
5.500 0.6035 0.04108 0.03168 0.0094 0.1379 1.0000
5.750 0.6256 0.04509 0.03619 0.0095 0.1394 1.0000
6.000 0.6477 0.05051 0.04224 0.0088 0.1546 1.0000
6.750 0.6477 0.08404 0.07745 -0.0462 0.4631 1.0000
7.000 0.6512 0.08665 0.08002 -0.0431 0.4251 1.0000
7.500 0.6619 0.09338 0.08668 -0.0396 0.3714 1.0000
7.750 0.6788 0.09793 0.09129 -0.0378 0.3508 1.0000
8.000 0.6707 0.10053 0.09381 -0.0378 0.3309 1.0000
8.250 0.6829 0.10461 0.09789 -0.0364 0.3121 1.0000
8.500 0.6919 0.10897 0.10223 -0.0354 0.2950 1.0000
8.750 0.6874 0.11200 0.10519 -0.0358 0.2795 1.0000
9.000 0.6867 0.11531 0.10845 -0.0360 0.2640 1.0000
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Polar data table (+)
Polar graphs
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