NASA/AMES A-01 AIRFOIL (ames01-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/AMES A-01 AIRFOIL (ames01-il) Reynolds number: 200,000 Max Cl/Cd: 55.32 at α=8.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ames01-il-200000-n5.txt Download as CSV file: xf-ames01-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.7277 0.06507 0.06155 -0.0217 1.0000 0.0207
-9.500 -0.7634 0.05562 0.05177 -0.0234 1.0000 0.0207
-9.250 -0.8094 0.04179 0.03688 -0.0222 1.0000 0.0213
-9.000 -0.7946 0.04004 0.03502 -0.0219 1.0000 0.0218
-8.750 -0.7755 0.03898 0.03386 -0.0217 1.0000 0.0224
-8.500 -0.7595 0.03688 0.03152 -0.0212 1.0000 0.0230
-8.250 -0.7419 0.03493 0.02934 -0.0208 1.0000 0.0238
-8.000 -0.7262 0.03211 0.02612 -0.0199 1.0000 0.0250
-7.750 -0.7108 0.02896 0.02242 -0.0187 1.0000 0.0261
-7.500 -0.6902 0.02653 0.01949 -0.0180 0.9942 0.0270
-7.250 -0.6636 0.02489 0.01769 -0.0185 0.9876 0.0278
-7.000 -0.6345 0.02403 0.01674 -0.0193 0.9828 0.0288
-6.750 -0.6060 0.02327 0.01586 -0.0199 0.9780 0.0301
-6.500 -0.5769 0.02226 0.01465 -0.0204 0.9739 0.0318
-6.250 -0.5490 0.02114 0.01330 -0.0205 0.9693 0.0332
-6.000 -0.5213 0.02020 0.01217 -0.0206 0.9646 0.0343
-5.750 -0.4944 0.01925 0.01120 -0.0207 0.9605 0.0359
-5.500 -0.4678 0.01880 0.01073 -0.0206 0.9557 0.0378
-5.250 -0.4412 0.01826 0.01012 -0.0204 0.9510 0.0398
-5.000 -0.4149 0.01767 0.00942 -0.0200 0.9468 0.0417
-4.750 -0.3886 0.01723 0.00889 -0.0197 0.9422 0.0434
-4.500 -0.3641 0.01650 0.00820 -0.0191 0.9373 0.0461
-4.250 -0.3389 0.01609 0.00779 -0.0186 0.9331 0.0486
-4.000 -0.3132 0.01567 0.00732 -0.0181 0.9284 0.0512
-3.750 -0.2875 0.01534 0.00693 -0.0176 0.9234 0.0539
-3.500 -0.2642 0.01489 0.00649 -0.0166 0.9182 0.0574
-3.250 -0.2397 0.01455 0.00615 -0.0158 0.9104 0.0612
-3.000 -0.2174 0.01428 0.00583 -0.0143 0.9029 0.0650
-2.750 -0.1932 0.01395 0.00548 -0.0133 0.8930 0.0700
-2.500 -0.1697 0.01363 0.00517 -0.0121 0.8846 0.0767
-2.250 -0.1454 0.01335 0.00485 -0.0111 0.8757 0.0844
-2.000 -0.1204 0.01300 0.00457 -0.0103 0.8668 0.1006
-1.750 -0.0974 0.01244 0.00427 -0.0092 0.8586 0.1605
-1.500 -0.0762 0.01133 0.00399 -0.0083 0.8479 0.3660
-1.250 -0.0569 0.01038 0.00380 -0.0064 0.8368 0.5605
-1.000 -0.0383 0.00967 0.00373 -0.0038 0.8247 0.7198
-0.750 -0.0143 0.00932 0.00378 -0.0017 0.8127 0.8419
-0.500 0.0260 0.00936 0.00391 -0.0032 0.8007 0.9174
-0.250 0.0663 0.00947 0.00397 -0.0050 0.7866 0.9550
0.000 0.1129 0.00958 0.00398 -0.0084 0.7677 0.9769
0.250 0.1538 0.00961 0.00389 -0.0109 0.7431 0.9876
0.500 0.1909 0.00963 0.00374 -0.0126 0.7095 0.9949
0.750 0.2269 0.00970 0.00353 -0.0141 0.6532 1.0000
1.000 0.2514 0.00992 0.00333 -0.0133 0.5735 1.0000
1.250 0.2769 0.01035 0.00326 -0.0131 0.4788 1.0000
1.500 0.3025 0.01082 0.00330 -0.0131 0.3997 1.0000
1.750 0.3280 0.01118 0.00336 -0.0129 0.3521 1.0000
2.000 0.3534 0.01146 0.00345 -0.0127 0.3242 1.0000
2.250 0.3787 0.01168 0.00354 -0.0124 0.3060 1.0000
2.500 0.4040 0.01190 0.00365 -0.0121 0.2923 1.0000
2.750 0.4292 0.01211 0.00377 -0.0117 0.2806 1.0000
3.000 0.4544 0.01231 0.00391 -0.0113 0.2711 1.0000
3.500 0.5048 0.01274 0.00424 -0.0105 0.2566 1.0000
3.750 0.5298 0.01298 0.00442 -0.0100 0.2504 1.0000
4.000 0.5549 0.01322 0.00462 -0.0096 0.2446 1.0000
4.250 0.5802 0.01344 0.00484 -0.0092 0.2386 1.0000
4.500 0.6050 0.01372 0.00507 -0.0087 0.2335 1.0000
4.750 0.6302 0.01397 0.00533 -0.0083 0.2291 1.0000
5.000 0.6554 0.01421 0.00559 -0.0078 0.2240 1.0000
5.250 0.6803 0.01451 0.00586 -0.0074 0.2191 1.0000
5.500 0.7053 0.01481 0.00616 -0.0070 0.2148 1.0000
5.750 0.7306 0.01506 0.00647 -0.0066 0.2099 1.0000
6.000 0.7556 0.01535 0.00677 -0.0062 0.2050 1.0000
6.250 0.7805 0.01569 0.00710 -0.0058 0.2003 1.0000
6.500 0.8060 0.01593 0.00743 -0.0054 0.1948 1.0000
6.750 0.8311 0.01622 0.00775 -0.0050 0.1896 1.0000
7.000 0.8561 0.01653 0.00809 -0.0047 0.1841 1.0000
7.250 0.8815 0.01677 0.00842 -0.0043 0.1776 1.0000
7.500 0.9061 0.01710 0.00875 -0.0040 0.1718 1.0000
7.750 0.9315 0.01735 0.00912 -0.0037 0.1648 1.0000
8.000 0.9560 0.01765 0.00943 -0.0033 0.1578 1.0000
8.250 0.9810 0.01791 0.00979 -0.0030 0.1488 1.0000
8.500 1.0055 0.01824 0.01017 -0.0026 0.1395 1.0000
8.750 1.0295 0.01861 0.01056 -0.0023 0.1295 1.0000
9.000 1.0529 0.01905 0.01101 -0.0018 0.1185 1.0000
9.250 1.0761 0.01954 0.01153 -0.0014 0.1062 1.0000
9.500 1.0982 0.02015 0.01215 -0.0009 0.0929 1.0000
9.750 1.1194 0.02087 0.01285 -0.0002 0.0798 1.0000
10.000 1.1395 0.02171 0.01368 0.0005 0.0685 1.0000
10.500 1.1773 0.02353 0.01557 0.0022 0.0514 1.0000
10.750 1.1947 0.02451 0.01662 0.0032 0.0458 1.0000
11.000 1.2109 0.02555 0.01773 0.0043 0.0413 1.0000
11.250 1.2249 0.02671 0.01896 0.0056 0.0380 1.0000
11.500 1.2392 0.02775 0.02014 0.0069 0.0352 1.0000
11.750 1.2504 0.02896 0.02144 0.0085 0.0328 1.0000
12.000 1.2549 0.03042 0.02297 0.0107 0.0309 1.0000
12.250 1.2610 0.03173 0.02444 0.0128 0.0296 1.0000
12.500 1.2663 0.03321 0.02606 0.0145 0.0283 1.0000
12.750 1.2698 0.03491 0.02790 0.0161 0.0272 1.0000
13.000 1.2718 0.03684 0.02996 0.0174 0.0263 1.0000
13.250 1.2716 0.03908 0.03232 0.0184 0.0255 1.0000
13.500 1.2685 0.04177 0.03512 0.0191 0.0248 1.0000
13.750 1.2618 0.04505 0.03853 0.0192 0.0242 1.0000
14.000 1.2542 0.04873 0.04233 0.0187 0.0237 1.0000
14.250 1.2491 0.05249 0.04628 0.0177 0.0233 1.0000
14.500 1.2414 0.05690 0.05087 0.0159 0.0229 1.0000
14.750 1.2316 0.06203 0.05617 0.0135 0.0226 1.0000
15.000 1.2198 0.06779 0.06212 0.0104 0.0223 1.0000
15.250 1.2064 0.07406 0.06854 0.0071 0.0220 1.0000
15.500 1.1920 0.08057 0.07522 0.0036 0.0218 1.0000
15.750 1.1774 0.08714 0.08192 0.0003 0.0216 1.0000
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