NASA/AMES A-03 AIRFOIL (ames03-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/AMES A-03 AIRFOIL (ames03-il) Reynolds number: 500,000 Max Cl/Cd: 72.71 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ames03-il-500000.txt Download as CSV file: xf-ames03-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES A-03 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5402 0.10452 0.10250 0.0182 0.8803 0.0245
-9.750 -0.5476 0.09846 0.09646 0.0154 0.8790 0.0253
-7.500 -0.7463 0.04000 0.03601 0.0010 0.8663 0.0206
-7.250 -0.7398 0.03215 0.02769 0.0015 0.8642 0.0187
-7.000 -0.7218 0.02705 0.02194 0.0024 0.8621 0.0184
-6.750 -0.6976 0.02487 0.01943 0.0027 0.8602 0.0190
-6.500 -0.6724 0.02302 0.01727 0.0030 0.8584 0.0197
-6.250 -0.6467 0.02137 0.01534 0.0034 0.8567 0.0201
-6.000 -0.6203 0.02032 0.01409 0.0038 0.8552 0.0205
-5.750 -0.5955 0.01824 0.01179 0.0042 0.8538 0.0213
-5.500 -0.5696 0.01719 0.01068 0.0045 0.8525 0.0222
-5.250 -0.5402 0.01652 0.00999 0.0040 0.8508 0.0233
-5.000 -0.5109 0.01598 0.00942 0.0035 0.8490 0.0248
-4.750 -0.4817 0.01555 0.00892 0.0032 0.8471 0.0262
-4.500 -0.4540 0.01457 0.00788 0.0031 0.8454 0.0277
-4.250 -0.4264 0.01394 0.00725 0.0030 0.8434 0.0298
-4.000 -0.3995 0.01353 0.00680 0.0034 0.8410 0.0322
-3.750 -0.3732 0.01321 0.00642 0.0039 0.8387 0.0346
-3.500 -0.3453 0.01260 0.00581 0.0038 0.8362 0.0396
-3.250 -0.3152 0.01232 0.00551 0.0033 0.8329 0.0439
-3.000 -0.2866 0.01185 0.00507 0.0032 0.8299 0.0543
-2.750 -0.2585 0.01140 0.00471 0.0031 0.8275 0.0763
-2.500 -0.2312 0.01084 0.00441 0.0031 0.8253 0.1393
-2.250 -0.2049 0.01007 0.00414 0.0031 0.8230 0.2686
-2.000 -0.1764 0.00886 0.00388 0.0021 0.8193 0.5091
-1.750 -0.1488 0.00813 0.00380 0.0020 0.8158 0.6804
-1.500 -0.1220 0.00786 0.00376 0.0025 0.8127 0.7520
-1.250 -0.0965 0.00769 0.00370 0.0035 0.8098 0.8017
-1.000 -0.0708 0.00760 0.00366 0.0045 0.8063 0.8365
-0.750 -0.0436 0.00750 0.00362 0.0050 0.8009 0.8600
-0.500 -0.0185 0.00739 0.00350 0.0061 0.7960 0.8813
-0.250 0.0047 0.00728 0.00339 0.0078 0.7903 0.9010
0.000 0.0282 0.00715 0.00328 0.0093 0.7830 0.9192
0.250 0.0476 0.00702 0.00312 0.0120 0.7777 0.9405
0.500 0.0690 0.00685 0.00299 0.0140 0.7706 0.9601
0.750 0.0955 0.00668 0.00278 0.0147 0.7640 0.9728
1.000 0.1284 0.00656 0.00264 0.0138 0.7542 0.9809
1.250 0.1638 0.00647 0.00251 0.0124 0.7431 0.9868
1.500 0.2002 0.00640 0.00240 0.0106 0.7284 0.9919
1.750 0.2372 0.00634 0.00228 0.0088 0.7040 0.9959
2.000 0.2748 0.00650 0.00209 0.0067 0.6050 0.9987
2.250 0.3169 0.00812 0.00249 0.0020 0.3391 0.9992
2.500 0.3471 0.00850 0.00263 0.0009 0.2936 1.0000
2.750 0.3730 0.00872 0.00272 0.0008 0.2741 1.0000
3.000 0.4003 0.00893 0.00284 0.0005 0.2609 1.0000
3.250 0.4290 0.00913 0.00298 0.0000 0.2509 1.0000
3.500 0.4583 0.00936 0.00315 -0.0007 0.2424 1.0000
3.750 0.4879 0.00955 0.00331 -0.0013 0.2351 1.0000
4.000 0.5176 0.00978 0.00349 -0.0020 0.2279 1.0000
4.250 0.5473 0.01000 0.00369 -0.0026 0.2215 1.0000
4.500 0.5768 0.01019 0.00387 -0.0032 0.2154 1.0000
4.750 0.6062 0.01050 0.00412 -0.0039 0.2090 1.0000
5.000 0.6357 0.01063 0.00430 -0.0044 0.2035 1.0000
5.250 0.6649 0.01088 0.00451 -0.0050 0.1972 1.0000
5.500 0.6941 0.01111 0.00475 -0.0055 0.1917 1.0000
5.750 0.7232 0.01128 0.00494 -0.0059 0.1858 1.0000
6.000 0.7519 0.01160 0.00522 -0.0065 0.1795 1.0000
6.250 0.7810 0.01173 0.00542 -0.0069 0.1742 1.0000
6.500 0.8096 0.01197 0.00563 -0.0073 0.1674 1.0000
6.750 0.8384 0.01215 0.00586 -0.0077 0.1604 1.0000
7.000 0.8667 0.01241 0.00607 -0.0082 0.1514 1.0000
7.250 0.8953 0.01256 0.00626 -0.0086 0.1428 1.0000
7.500 0.9233 0.01283 0.00653 -0.0089 0.1340 1.0000
7.750 0.9510 0.01316 0.00683 -0.0093 0.1236 1.0000
8.000 0.9787 0.01346 0.00713 -0.0096 0.1131 1.0000
8.250 1.0059 0.01384 0.00749 -0.0099 0.1016 1.0000
8.500 1.0325 0.01429 0.00791 -0.0102 0.0891 1.0000
8.750 1.0586 0.01482 0.00840 -0.0104 0.0769 1.0000
9.000 1.0842 0.01542 0.00895 -0.0106 0.0650 1.0000
9.250 1.1089 0.01613 0.00961 -0.0106 0.0527 1.0000
9.500 1.1328 0.01697 0.01040 -0.0106 0.0416 1.0000
9.750 1.1567 0.01775 0.01119 -0.0106 0.0346 1.0000
10.000 1.1797 0.01863 0.01212 -0.0104 0.0301 1.0000
10.250 1.2019 0.01954 0.01305 -0.0102 0.0269 1.0000
10.500 1.2223 0.02063 0.01424 -0.0098 0.0245 1.0000
10.750 1.2442 0.02142 0.01511 -0.0095 0.0228 1.0000
11.000 1.2643 0.02235 0.01612 -0.0091 0.0213 1.0000
11.250 1.2790 0.02383 0.01768 -0.0083 0.0199 1.0000
11.500 1.2925 0.02529 0.01927 -0.0073 0.0191 1.0000
11.750 1.3084 0.02638 0.02049 -0.0066 0.0185 1.0000
12.000 1.3197 0.02761 0.02184 -0.0054 0.0179 1.0000
12.250 1.3282 0.02898 0.02331 -0.0039 0.0173 1.0000
12.500 1.3360 0.03046 0.02490 -0.0025 0.0168 1.0000
12.750 1.3424 0.03210 0.02663 -0.0012 0.0163 1.0000
13.000 1.3468 0.03397 0.02862 0.0000 0.0159 1.0000
13.250 1.3477 0.03621 0.03097 0.0012 0.0156 1.0000
13.500 1.3438 0.03902 0.03391 0.0024 0.0152 1.0000
13.750 1.3351 0.04247 0.03751 0.0034 0.0149 1.0000
14.000 1.3249 0.04627 0.04149 0.0041 0.0147 1.0000
14.250 1.3242 0.04915 0.04452 0.0040 0.0146 1.0000
14.500 1.3210 0.05250 0.04803 0.0035 0.0144 1.0000
14.750 1.3155 0.05634 0.05203 0.0026 0.0142 1.0000
15.000 1.3078 0.06082 0.05668 0.0010 0.0141 1.0000
15.250 1.2975 0.06608 0.06210 -0.0014 0.0139 1.0000
15.500 1.2856 0.07218 0.06838 -0.0048 0.0138 1.0000
15.750 1.2718 0.07907 0.07543 -0.0088 0.0138 1.0000
16.000 1.2561 0.08654 0.08307 -0.0131 0.0137 1.0000
16.250 1.2387 0.09442 0.09111 -0.0176 0.0137 1.0000
16.500 1.2198 0.10254 0.09937 -0.0222 0.0137 1.0000
16.750 1.2001 0.11085 0.10782 -0.0268 0.0137 1.0000
17.000 1.1797 0.11939 0.11650 -0.0316 0.0138 1.0000
17.250 1.1589 0.12815 0.12538 -0.0365 0.0138 1.0000
17.500 1.1375 0.13748 0.13484 -0.0419 0.0139 1.0000
17.750 1.1142 0.14769 0.14519 -0.0479 0.0139 1.0000
18.000 1.0860 0.15976 0.15740 -0.0550 0.0141 1.0000
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