AG08 (ag08-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: AG08 (ag08-il) Reynolds number: 50,000 Max Cl/Cd: 32.76 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ag08-il-50000.txt Download as CSV file: xf-ag08-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: AG08
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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.5646 0.10478 0.09829 0.0159 1.0000 0.1978
-7.500 -0.5510 0.09987 0.09334 0.0170 1.0000 0.2060
-7.250 -0.5576 0.09759 0.09119 0.0122 1.0000 0.2138
-7.000 -0.5498 0.09404 0.08767 0.0110 1.0000 0.2260
-6.750 -0.5391 0.08985 0.08351 0.0115 1.0000 0.2389
-6.500 -0.5350 0.08677 0.08048 0.0075 1.0000 0.2544
-6.000 -0.5150 0.07927 0.07306 0.0087 1.0000 0.2959
-5.750 -0.5030 0.07510 0.06893 0.0106 1.0000 0.3146
-5.500 -0.4943 0.07179 0.06566 0.0095 1.0000 0.3404
-5.250 -0.4836 0.06831 0.06222 0.0107 1.0000 0.3684
-5.000 -0.4719 0.06473 0.05868 0.0135 1.0000 0.3983
-4.750 -0.3807 0.04724 0.03945 -0.0340 1.0000 0.1571
-4.500 -0.3523 0.04155 0.03336 -0.0361 1.0000 0.1417
-4.250 -0.3244 0.03731 0.02864 -0.0375 1.0000 0.1412
-4.000 -0.2956 0.03328 0.02410 -0.0384 1.0000 0.1388
-3.750 -0.2655 0.02982 0.01987 -0.0389 1.0000 0.1412
-3.500 -0.2380 0.02743 0.01711 -0.0387 1.0000 0.1554
-3.250 -0.2100 0.02497 0.01436 -0.0383 1.0000 0.1666
-3.000 -0.1833 0.02319 0.01245 -0.0378 1.0000 0.1926
-2.750 -0.1557 0.02138 0.01051 -0.0371 1.0000 0.2212
-2.500 -0.1293 0.01991 0.00906 -0.0362 1.0000 0.2698
-2.250 -0.1024 0.01836 0.00781 -0.0354 1.0000 0.3490
-2.000 -0.0806 0.01595 0.00662 -0.0335 1.0000 0.5387
-1.750 -0.0454 0.01403 0.00536 -0.0326 1.0000 1.0000
-1.500 -0.0196 0.01402 0.00488 -0.0321 1.0000 1.0000
-1.250 0.0056 0.01402 0.00451 -0.0317 1.0000 1.0000
-1.000 0.0304 0.01404 0.00426 -0.0311 1.0000 1.0000
-0.750 0.0550 0.01408 0.00409 -0.0306 1.0000 1.0000
-0.500 0.0793 0.01414 0.00398 -0.0301 1.0000 1.0000
-0.250 0.1036 0.01422 0.00392 -0.0296 1.0000 1.0000
0.000 0.1277 0.01432 0.00389 -0.0291 1.0000 1.0000
0.250 0.1515 0.01445 0.00395 -0.0286 1.0000 1.0000
0.500 0.1751 0.01461 0.00406 -0.0282 1.0000 1.0000
0.750 0.1984 0.01480 0.00424 -0.0278 1.0000 1.0000
1.000 0.2215 0.01504 0.00448 -0.0275 1.0000 1.0000
1.250 0.2443 0.01534 0.00480 -0.0273 1.0000 1.0000
1.500 0.2666 0.01571 0.00522 -0.0272 1.0000 1.0000
1.750 0.2881 0.01619 0.00576 -0.0273 1.0000 1.0000
2.000 0.3085 0.01682 0.00647 -0.0277 1.0000 1.0000
2.250 0.3273 0.01765 0.00744 -0.0285 1.0000 1.0000
2.500 0.4004 0.01859 0.00866 -0.0394 0.9685 1.0000
2.750 0.4958 0.01863 0.00920 -0.0514 0.9083 1.0000
3.000 0.5511 0.01837 0.00920 -0.0538 0.8474 1.0000
3.250 0.5814 0.01834 0.00921 -0.0510 0.7885 1.0000
3.500 0.6029 0.01862 0.00939 -0.0472 0.7298 1.0000
3.750 0.6235 0.01908 0.00978 -0.0437 0.6726 1.0000
4.000 0.6447 0.01968 0.01022 -0.0407 0.6177 1.0000
4.250 0.6667 0.02040 0.01077 -0.0381 0.5652 1.0000
4.500 0.6892 0.02126 0.01151 -0.0361 0.5134 1.0000
4.750 0.7119 0.02220 0.01230 -0.0341 0.4649 1.0000
5.000 0.7345 0.02324 0.01331 -0.0324 0.4162 1.0000
5.250 0.7571 0.02437 0.01425 -0.0307 0.3699 1.0000
5.500 0.7795 0.02566 0.01553 -0.0292 0.3221 1.0000
5.750 0.8015 0.02703 0.01678 -0.0277 0.2759 1.0000
6.000 0.8235 0.02864 0.01824 -0.0262 0.2325 1.0000
6.250 0.8457 0.03068 0.02040 -0.0249 0.1940 1.0000
6.500 0.8677 0.03286 0.02238 -0.0237 0.1627 1.0000
6.750 0.8893 0.03544 0.02535 -0.0225 0.1388 1.0000
7.000 0.9110 0.03879 0.02902 -0.0214 0.1248 1.0000
7.250 0.9290 0.04271 0.03355 -0.0205 0.1145 1.0000
7.500 0.9457 0.04662 0.03779 -0.0197 0.1065 1.0000
7.750 0.9566 0.05216 0.04405 -0.0192 0.1052 1.0000
8.000 0.9624 0.05822 0.05071 -0.0192 0.1055 1.0000
8.250 0.9640 0.06447 0.05740 -0.0198 0.1064 1.0000
8.500 0.9626 0.07072 0.06396 -0.0208 0.1074 1.0000
8.750 0.9605 0.07686 0.07028 -0.0220 0.1083 1.0000
9.000 0.9590 0.08292 0.07649 -0.0232 0.1094 1.0000
9.250 0.8981 0.09869 0.09236 -0.0404 0.1330 1.0000
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Polar data table (+)
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