NASA/AMES A-03 AIRFOIL (ames03-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/AMES A-03 AIRFOIL (ames03-il) Reynolds number: 100,000 Max Cl/Cd: 39.56 at α=8.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ames03-il-100000-n5.txt Download as CSV file: xf-ames03-il-100000-n5.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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6607 0.06437 0.05953 -0.0299 0.9762 0.0266
-8.750 -0.6639 0.05937 0.05425 -0.0314 0.9693 0.0266
-8.500 -0.6641 0.05456 0.04907 -0.0321 0.9633 0.0266
-8.250 -0.6612 0.04982 0.04385 -0.0323 0.9583 0.0267
-8.000 -0.6504 0.04660 0.04043 -0.0324 0.9542 0.0274
-7.750 -0.6329 0.04504 0.03879 -0.0325 0.9506 0.0286
-7.500 -0.6174 0.04239 0.03583 -0.0324 0.9469 0.0298
-7.250 -0.6013 0.03910 0.03207 -0.0320 0.9429 0.0306
-7.000 -0.5834 0.03593 0.02838 -0.0314 0.9394 0.0311
-6.750 -0.5630 0.03323 0.02518 -0.0308 0.9365 0.0318
-6.500 -0.5390 0.03079 0.02224 -0.0306 0.9338 0.0328
-6.250 -0.5138 0.02895 0.01994 -0.0303 0.9310 0.0346
-6.000 -0.4893 0.02754 0.01848 -0.0304 0.9281 0.0367
-5.750 -0.4639 0.02628 0.01705 -0.0301 0.9254 0.0387
-5.500 -0.4382 0.02502 0.01560 -0.0297 0.9230 0.0406
-5.250 -0.4107 0.02389 0.01426 -0.0297 0.9205 0.0431
-5.000 -0.3847 0.02283 0.01321 -0.0298 0.9181 0.0470
-4.750 -0.3586 0.02201 0.01235 -0.0297 0.9157 0.0509
-4.500 -0.3330 0.02129 0.01150 -0.0294 0.9132 0.0552
-4.250 -0.3088 0.02058 0.01084 -0.0290 0.9108 0.0623
-4.000 -0.2831 0.01995 0.01016 -0.0288 0.9084 0.0706
-3.750 -0.2564 0.01934 0.00959 -0.0290 0.9060 0.0854
-3.500 -0.2301 0.01873 0.00909 -0.0290 0.9037 0.1113
-3.250 -0.2051 0.01795 0.00868 -0.0291 0.9014 0.1811
-3.000 -0.1835 0.01675 0.00839 -0.0287 0.8989 0.3659
-2.750 -0.1663 0.01576 0.00854 -0.0263 0.8967 0.6226
-2.500 -0.1449 0.01573 0.00886 -0.0239 0.8945 0.7442
-2.250 -0.1217 0.01590 0.00909 -0.0220 0.8922 0.8090
-2.000 -0.0986 0.01613 0.00930 -0.0201 0.8888 0.8547
-1.750 -0.0744 0.01636 0.00947 -0.0181 0.8844 0.8904
-1.500 -0.0453 0.01657 0.00959 -0.0169 0.8791 0.9239
-1.250 -0.0058 0.01666 0.00957 -0.0181 0.8721 0.9482
-1.000 0.0287 0.01663 0.00941 -0.0189 0.8661 0.9602
-0.750 0.0651 0.01656 0.00925 -0.0207 0.8590 0.9692
-0.500 0.1005 0.01645 0.00904 -0.0219 0.8530 0.9762
-0.250 0.1367 0.01635 0.00887 -0.0237 0.8446 0.9841
0.000 0.1697 0.01615 0.00860 -0.0243 0.8376 0.9912
0.250 0.2055 0.01597 0.00839 -0.0260 0.8272 0.9983
0.500 0.2275 0.01576 0.00814 -0.0247 0.8172 1.0000
0.750 0.2436 0.01546 0.00779 -0.0219 0.8072 1.0000
1.000 0.2634 0.01520 0.00751 -0.0201 0.7942 1.0000
1.250 0.2837 0.01496 0.00725 -0.0185 0.7809 1.0000
1.500 0.3044 0.01471 0.00700 -0.0168 0.7668 1.0000
1.750 0.3256 0.01447 0.00674 -0.0152 0.7508 1.0000
2.000 0.3480 0.01425 0.00653 -0.0140 0.7304 1.0000
2.250 0.3718 0.01408 0.00638 -0.0131 0.7013 1.0000
2.500 0.3943 0.01385 0.00611 -0.0116 0.6462 1.0000
2.750 0.4040 0.01353 0.00492 -0.0058 0.4962 1.0000
3.000 0.4263 0.01444 0.00512 -0.0053 0.3876 1.0000
3.250 0.4512 0.01505 0.00543 -0.0052 0.3430 1.0000
3.500 0.4765 0.01553 0.00572 -0.0052 0.3182 1.0000
3.750 0.5020 0.01597 0.00602 -0.0050 0.3010 1.0000
4.000 0.5278 0.01637 0.00633 -0.0049 0.2873 1.0000
4.250 0.5538 0.01676 0.00668 -0.0048 0.2763 1.0000
4.500 0.5797 0.01719 0.00704 -0.0046 0.2669 1.0000
4.750 0.6057 0.01762 0.00743 -0.0045 0.2577 1.0000
5.000 0.6320 0.01806 0.00785 -0.0044 0.2493 1.0000
5.250 0.6580 0.01854 0.00832 -0.0043 0.2416 1.0000
5.500 0.6844 0.01900 0.00881 -0.0043 0.2335 1.0000
5.750 0.7105 0.01951 0.00930 -0.0043 0.2262 1.0000
6.000 0.7369 0.01999 0.00988 -0.0043 0.2184 1.0000
6.250 0.7629 0.02052 0.01040 -0.0042 0.2110 1.0000
6.500 0.7893 0.02102 0.01102 -0.0042 0.2032 1.0000
6.750 0.8150 0.02155 0.01156 -0.0042 0.1957 1.0000
7.000 0.8411 0.02205 0.01225 -0.0043 0.1874 1.0000
7.250 0.8665 0.02258 0.01278 -0.0042 0.1800 1.0000
7.500 0.8922 0.02307 0.01348 -0.0042 0.1708 1.0000
7.750 0.9172 0.02357 0.01408 -0.0042 0.1625 1.0000
8.000 0.9420 0.02405 0.01470 -0.0041 0.1533 1.0000
8.250 0.9664 0.02456 0.01538 -0.0041 0.1429 1.0000
8.500 0.9902 0.02507 0.01600 -0.0040 0.1324 1.0000
8.750 1.0132 0.02561 0.01663 -0.0038 0.1211 1.0000
9.000 1.0356 0.02624 0.01736 -0.0036 0.1094 1.0000
9.250 1.0570 0.02700 0.01818 -0.0034 0.0976 1.0000
9.500 1.0774 0.02792 0.01918 -0.0031 0.0859 1.0000
9.750 1.0968 0.02902 0.02039 -0.0026 0.0753 1.0000
10.000 1.1145 0.03031 0.02180 -0.0020 0.0664 1.0000
10.250 1.1295 0.03179 0.02336 -0.0013 0.0595 1.0000
10.500 1.1438 0.03327 0.02499 -0.0005 0.0532 1.0000
10.750 1.1544 0.03503 0.02685 0.0006 0.0489 1.0000
11.000 1.1646 0.03681 0.02884 0.0017 0.0453 1.0000
11.250 1.1700 0.03860 0.03075 0.0030 0.0426 1.0000
11.500 1.1714 0.04074 0.03295 0.0045 0.0405 1.0000
11.750 1.1752 0.04291 0.03540 0.0058 0.0383 1.0000
12.000 1.1765 0.04529 0.03798 0.0068 0.0364 1.0000
12.250 1.1759 0.04788 0.04077 0.0077 0.0351 1.0000
12.500 1.1736 0.05072 0.04376 0.0083 0.0341 1.0000
12.750 1.1694 0.05388 0.04705 0.0085 0.0333 1.0000
13.000 1.1635 0.05740 0.05068 0.0084 0.0326 1.0000
13.250 1.1550 0.06153 0.05501 0.0078 0.0320 1.0000
13.500 1.1432 0.06644 0.06018 0.0063 0.0315 1.0000
13.750 1.1285 0.07221 0.06621 0.0037 0.0312 1.0000
14.000 1.1110 0.07912 0.07335 -0.0002 0.0310 1.0000
14.250 1.0907 0.08731 0.08177 -0.0053 0.0310 1.0000
14.500 1.0676 0.09676 0.09141 -0.0114 0.0311 1.0000
14.750 1.0415 0.10721 0.10203 -0.0178 0.0314 1.0000
15.000 1.0115 0.11879 0.11372 -0.0247 0.0317 1.0000
15.250 0.9769 0.13194 0.12697 -0.0322 0.0321 1.0000
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Polar data table (+)
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