NASA/AMES A-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il) Reynolds number: 50,000 Max Cl/Cd: 24.47 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ames02-il-50000.txt Download as CSV file: xf-ames02-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES A-02 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
------ -------- --------- --------- -------- -------- --------
-7.250 -0.6232 0.09718 0.09054 0.0417 1.0000 0.3521
-7.000 -0.6190 0.09400 0.08741 0.0425 1.0000 0.3739
-6.750 -0.6391 0.09243 0.08599 0.0426 1.0000 0.3979
-6.500 -0.5982 0.08721 0.08070 0.0459 1.0000 0.4271
-6.250 -0.5932 0.08416 0.07772 0.0479 1.0000 0.4573
-5.750 -0.6165 0.05180 0.04345 -0.0104 1.0000 0.1474
-5.500 -0.5933 0.04639 0.03726 -0.0118 1.0000 0.1337
-5.250 -0.5695 0.04239 0.03298 -0.0120 1.0000 0.1317
-5.000 -0.5441 0.03869 0.02884 -0.0122 1.0000 0.1293
-4.750 -0.5169 0.03529 0.02489 -0.0122 1.0000 0.1270
-4.500 -0.4890 0.03261 0.02176 -0.0119 1.0000 0.1297
-4.250 -0.4599 0.03031 0.01890 -0.0116 1.0000 0.1338
-4.000 -0.4321 0.02792 0.01650 -0.0112 1.0000 0.1404
-3.750 -0.4032 0.02599 0.01434 -0.0107 1.0000 0.1507
-3.500 -0.3750 0.02418 0.01256 -0.0100 1.0000 0.1662
-3.250 -0.3474 0.02242 0.01098 -0.0092 1.0000 0.1911
-3.000 -0.3209 0.02045 0.00937 -0.0084 1.0000 0.2422
-2.750 -0.3185 0.01656 0.00863 -0.0010 1.0000 0.6988
-2.500 -0.1878 0.01757 0.00895 -0.0095 1.0000 1.0000
-2.250 -0.1680 0.01699 0.00813 -0.0097 1.0000 1.0000
-2.000 -0.1482 0.01650 0.00745 -0.0096 1.0000 1.0000
-1.750 -0.1286 0.01610 0.00689 -0.0093 1.0000 1.0000
-1.500 -0.1095 0.01579 0.00642 -0.0087 1.0000 1.0000
-1.250 -0.0904 0.01554 0.00605 -0.0079 1.0000 1.0000
-1.000 -0.0714 0.01537 0.00574 -0.0070 1.0000 1.0000
-0.750 -0.0525 0.01526 0.00552 -0.0060 1.0000 1.0000
-0.500 -0.0336 0.01521 0.00538 -0.0049 1.0000 1.0000
-0.250 -0.0148 0.01521 0.00529 -0.0038 1.0000 1.0000
0.000 0.0041 0.01525 0.00527 -0.0026 1.0000 1.0000
0.250 0.0231 0.01534 0.00530 -0.0015 1.0000 1.0000
0.500 0.0426 0.01546 0.00539 -0.0004 1.0000 1.0000
0.750 0.0625 0.01562 0.00553 0.0005 1.0000 1.0000
1.000 0.0827 0.01581 0.00573 0.0013 1.0000 1.0000
1.250 0.1034 0.01605 0.00598 0.0020 1.0000 1.0000
1.500 0.1243 0.01634 0.00630 0.0026 1.0000 1.0000
1.750 0.1454 0.01667 0.00670 0.0030 1.0000 1.0000
2.000 0.1664 0.01706 0.00716 0.0033 1.0000 1.0000
2.250 0.1872 0.01751 0.00772 0.0034 1.0000 1.0000
2.500 0.2076 0.01806 0.00840 0.0034 1.0000 1.0000
2.750 0.2273 0.01873 0.00923 0.0031 1.0000 1.0000
3.000 0.3694 0.01876 0.01004 -0.0173 0.9284 1.0000
3.250 0.4204 0.01800 0.00966 -0.0160 0.8232 1.0000
3.500 0.4299 0.01757 0.00892 -0.0067 0.6729 1.0000
3.750 0.4393 0.01889 0.00880 0.0002 0.4763 1.0000
4.000 0.4608 0.02075 0.00982 0.0021 0.3697 1.0000
4.250 0.4857 0.02240 0.01104 0.0031 0.3100 1.0000
4.500 0.5117 0.02405 0.01242 0.0039 0.2702 1.0000
4.750 0.5387 0.02568 0.01406 0.0044 0.2395 1.0000
5.000 0.5653 0.02747 0.01574 0.0048 0.2171 1.0000
5.250 0.5920 0.02943 0.01772 0.0052 0.1991 1.0000
5.500 0.6191 0.03145 0.02004 0.0054 0.1844 1.0000
5.750 0.6458 0.03397 0.02299 0.0055 0.1731 1.0000
6.000 0.6710 0.03664 0.02577 0.0057 0.1645 1.0000
6.250 0.6955 0.03983 0.02961 0.0055 0.1577 1.0000
6.500 0.7189 0.04278 0.03263 0.0057 0.1521 1.0000
6.750 0.7389 0.04710 0.03757 0.0052 0.1496 1.0000
7.000 0.7566 0.05191 0.04289 0.0046 0.1498 1.0000
7.250 0.7719 0.05697 0.04834 0.0038 0.1504 1.0000
7.500 0.6704 0.09241 0.08519 -0.0478 0.4039 1.0000
7.750 0.6785 0.09602 0.08881 -0.0468 0.3845 1.0000
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Polar data table (+)
Polar graphs
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