NASA/AMES A-01 AIRFOIL (ames01-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/AMES A-01 AIRFOIL (ames01-il) Reynolds number: 100,000 Max Cl/Cd: 40.69 at α=9.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ames01-il-100000-n5.txt Download as CSV file: xf-ames01-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES A-01 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.6057 0.09616 0.09132 -0.0024 1.0000 0.0372
-9.500 -0.6113 0.08977 0.08496 -0.0071 1.0000 0.0373
-9.250 -0.6756 0.07020 0.06513 -0.0239 1.0000 0.0344
-9.000 -0.6686 0.06831 0.06327 -0.0238 1.0000 0.0352
-8.750 -0.6712 0.06427 0.05910 -0.0245 1.0000 0.0356
-8.500 -0.6744 0.05952 0.05413 -0.0250 1.0000 0.0358
-8.250 -0.6752 0.05462 0.04891 -0.0250 1.0000 0.0359
-8.000 -0.6707 0.05038 0.04436 -0.0245 1.0000 0.0362
-7.750 -0.6624 0.04665 0.04028 -0.0239 1.0000 0.0367
-7.500 -0.6511 0.04342 0.03672 -0.0230 1.0000 0.0375
-7.250 -0.6377 0.04072 0.03370 -0.0220 1.0000 0.0390
-7.000 -0.6245 0.03788 0.03042 -0.0206 1.0000 0.0406
-6.750 -0.6114 0.03516 0.02722 -0.0187 1.0000 0.0416
-6.500 -0.5980 0.03283 0.02444 -0.0166 1.0000 0.0425
-6.250 -0.5839 0.03101 0.02218 -0.0143 1.0000 0.0435
-6.000 -0.5688 0.02937 0.02047 -0.0127 1.0000 0.0450
-5.750 -0.5523 0.02837 0.01939 -0.0111 1.0000 0.0470
-5.500 -0.5349 0.02725 0.01810 -0.0094 1.0000 0.0489
-5.250 -0.5163 0.02599 0.01662 -0.0078 1.0000 0.0505
-5.000 -0.4903 0.02479 0.01516 -0.0075 0.9979 0.0529
-4.750 -0.4582 0.02355 0.01373 -0.0086 0.9945 0.0560
-4.500 -0.4275 0.02245 0.01264 -0.0095 0.9907 0.0588
-4.250 -0.3949 0.02151 0.01165 -0.0106 0.9871 0.0618
-4.000 -0.3612 0.02081 0.01082 -0.0119 0.9841 0.0665
-3.750 -0.3309 0.01992 0.00996 -0.0127 0.9798 0.0706
-3.500 -0.2986 0.01923 0.00929 -0.0138 0.9761 0.0752
-3.250 -0.2637 0.01874 0.00872 -0.0154 0.9731 0.0821
-3.000 -0.2340 0.01812 0.00815 -0.0161 0.9682 0.0890
-2.750 -0.2018 0.01768 0.00769 -0.0171 0.9636 0.0981
-2.500 -0.1677 0.01718 0.00728 -0.0186 0.9599 0.1156
-2.250 -0.1382 0.01655 0.00690 -0.0191 0.9524 0.1594
-2.000 -0.1100 0.01475 0.00665 -0.0198 0.9450 0.4925
-1.750 -0.0833 0.01394 0.00682 -0.0182 0.9343 0.7322
-1.500 -0.0361 0.01394 0.00714 -0.0195 0.9230 0.8776
-1.250 0.0210 0.01423 0.00736 -0.0234 0.9130 0.9489
-1.000 0.0836 0.01434 0.00732 -0.0292 0.9023 0.9834
-0.750 0.1258 0.01420 0.00704 -0.0317 0.8861 0.9959
-0.500 0.1531 0.01403 0.00678 -0.0315 0.8701 1.0000
-0.250 0.1734 0.01389 0.00655 -0.0299 0.8557 1.0000
0.000 0.1941 0.01376 0.00634 -0.0283 0.8411 1.0000
0.250 0.2153 0.01363 0.00615 -0.0268 0.8252 1.0000
0.500 0.2370 0.01352 0.00597 -0.0254 0.8072 1.0000
0.750 0.2587 0.01340 0.00579 -0.0239 0.7864 1.0000
1.000 0.2804 0.01329 0.00560 -0.0223 0.7614 1.0000
1.250 0.3021 0.01318 0.00540 -0.0207 0.7292 1.0000
1.500 0.3233 0.01309 0.00516 -0.0187 0.6848 1.0000
1.750 0.3436 0.01311 0.00485 -0.0165 0.6176 1.0000
2.000 0.3640 0.01343 0.00465 -0.0145 0.5249 1.0000
2.250 0.3859 0.01398 0.00469 -0.0134 0.4416 1.0000
2.500 0.4090 0.01449 0.00484 -0.0127 0.3903 1.0000
2.750 0.4326 0.01492 0.00504 -0.0120 0.3594 1.0000
3.000 0.4565 0.01531 0.00525 -0.0114 0.3388 1.0000
3.250 0.4806 0.01568 0.00548 -0.0108 0.3234 1.0000
3.500 0.5048 0.01603 0.00575 -0.0102 0.3112 1.0000
3.750 0.5291 0.01638 0.00603 -0.0097 0.3005 1.0000
4.000 0.5533 0.01676 0.00632 -0.0091 0.2916 1.0000
4.250 0.5778 0.01712 0.00665 -0.0085 0.2830 1.0000
4.500 0.6022 0.01753 0.00700 -0.0079 0.2764 1.0000
4.750 0.6269 0.01790 0.00738 -0.0074 0.2690 1.0000
5.000 0.6512 0.01836 0.00776 -0.0069 0.2629 1.0000
5.250 0.6763 0.01874 0.00821 -0.0064 0.2561 1.0000
5.500 0.7010 0.01919 0.00864 -0.0060 0.2504 1.0000
5.750 0.7259 0.01966 0.00912 -0.0055 0.2448 1.0000
6.000 0.7508 0.02010 0.00965 -0.0051 0.2384 1.0000
6.250 0.7755 0.02061 0.01012 -0.0047 0.2332 1.0000
6.500 0.8004 0.02109 0.01072 -0.0043 0.2272 1.0000
6.750 0.8251 0.02156 0.01127 -0.0039 0.2209 1.0000
7.000 0.8496 0.02210 0.01183 -0.0035 0.2155 1.0000
7.250 0.8740 0.02257 0.01247 -0.0030 0.2084 1.0000
7.500 0.8981 0.02307 0.01298 -0.0026 0.2026 1.0000
7.750 0.9220 0.02356 0.01367 -0.0021 0.1951 1.0000
8.000 0.9455 0.02398 0.01414 -0.0017 0.1882 1.0000
8.250 0.9687 0.02445 0.01481 -0.0011 0.1801 1.0000
8.500 0.9913 0.02482 0.01523 -0.0006 0.1728 1.0000
8.750 1.0139 0.02527 0.01592 0.0000 0.1635 1.0000
9.000 1.0357 0.02565 0.01643 0.0007 0.1543 1.0000
9.250 1.0567 0.02601 0.01685 0.0014 0.1447 1.0000
9.500 1.0776 0.02648 0.01751 0.0021 0.1330 1.0000
9.750 1.0976 0.02706 0.01821 0.0028 0.1207 1.0000
10.000 1.1165 0.02778 0.01902 0.0037 0.1082 1.0000
10.250 1.1338 0.02867 0.01997 0.0047 0.0959 1.0000
10.500 1.1493 0.02977 0.02113 0.0058 0.0849 1.0000
10.750 1.1625 0.03107 0.02251 0.0071 0.0759 1.0000
11.000 1.1722 0.03257 0.02401 0.0086 0.0687 1.0000
11.250 1.1816 0.03408 0.02569 0.0102 0.0624 1.0000
11.500 1.1849 0.03586 0.02749 0.0123 0.0584 1.0000
11.750 1.1878 0.03755 0.02938 0.0145 0.0547 1.0000
12.000 1.1880 0.03947 0.03144 0.0165 0.0517 1.0000
12.250 1.1849 0.04174 0.03379 0.0181 0.0494 1.0000
12.500 1.1811 0.04430 0.03646 0.0193 0.0475 1.0000
12.750 1.1782 0.04702 0.03940 0.0201 0.0455 1.0000
13.000 1.1730 0.05014 0.04270 0.0204 0.0439 1.0000
13.250 1.1660 0.05369 0.04645 0.0200 0.0427 1.0000
13.500 1.1574 0.05775 0.05066 0.0190 0.0417 1.0000
13.750 1.1471 0.06236 0.05542 0.0173 0.0410 1.0000
14.000 1.1359 0.06747 0.06066 0.0150 0.0403 1.0000
14.250 1.1245 0.07285 0.06615 0.0125 0.0397 1.0000
14.500 1.1151 0.07791 0.07126 0.0103 0.0390 1.0000
14.750 1.1003 0.08450 0.07806 0.0068 0.0386 1.0000
15.000 1.0839 0.09160 0.08535 0.0030 0.0383 1.0000
15.250 1.0656 0.09924 0.09317 -0.0012 0.0380 1.0000
15.500 1.0447 0.10760 0.10170 -0.0057 0.0378 1.0000
15.750 1.0207 0.11688 0.11115 -0.0108 0.0377 1.0000
16.000 0.9910 0.12786 0.12227 -0.0168 0.0377 1.0000
16.250 0.9414 0.14489 0.13944 -0.0261 0.0380 1.0000
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