AG10 (ag10-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: AG10 (ag10-il) Reynolds number: 50,000 Max Cl/Cd: 31.19 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ag10-il-50000.txt Download as CSV file: xf-ag10-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: AG10
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 -0.6084 0.10979 0.10309 0.0224 1.0000 0.1640
-7.500 -0.5904 0.10361 0.09691 0.0258 1.0000 0.1705
-7.250 -0.5916 0.10152 0.09489 0.0201 1.0000 0.1785
-6.750 -0.5756 0.09357 0.08703 0.0155 1.0000 0.1958
-6.500 -0.5660 0.09046 0.08391 0.0119 1.0000 0.2092
-6.250 -0.5531 0.08564 0.07914 0.0153 1.0000 0.2213
-6.000 -0.5418 0.08167 0.07521 0.0145 1.0000 0.2340
-5.750 -0.5301 0.07798 0.07151 0.0125 1.0000 0.2503
-5.500 -0.5180 0.07443 0.06797 0.0085 1.0000 0.2705
-5.250 -0.5050 0.07043 0.06401 0.0108 1.0000 0.2894
-5.000 -0.4925 0.06691 0.06052 0.0105 1.0000 0.3176
-4.750 -0.4812 0.06361 0.05728 0.0130 1.0000 0.3533
-4.500 -0.4737 0.06060 0.05435 0.0161 1.0000 0.4047
-4.250 -0.4667 0.05739 0.05123 0.0235 1.0000 0.4552
-4.000 -0.4618 0.05451 0.04846 0.0314 1.0000 0.5178
-3.500 -0.2715 0.03700 0.02808 -0.0300 1.0000 0.1627
-3.250 -0.2352 0.03309 0.02338 -0.0318 1.0000 0.1489
-3.000 -0.2046 0.02992 0.01983 -0.0323 1.0000 0.1471
-2.750 -0.1746 0.02732 0.01687 -0.0325 1.0000 0.1531
-2.500 -0.1444 0.02511 0.01426 -0.0324 1.0000 0.1601
-2.250 -0.1152 0.02321 0.01211 -0.0321 1.0000 0.1743
-2.000 -0.0862 0.02159 0.01021 -0.0316 1.0000 0.1949
-1.750 -0.0585 0.02005 0.00866 -0.0308 1.0000 0.2249
-1.500 -0.0304 0.01852 0.00730 -0.0301 1.0000 0.2740
-1.250 -0.0042 0.01655 0.00599 -0.0293 1.0000 0.3845
-1.000 0.0235 0.01377 0.00459 -0.0270 1.0000 1.0000
-0.750 0.0501 0.01377 0.00418 -0.0264 1.0000 1.0000
-0.500 0.0762 0.01379 0.00392 -0.0260 1.0000 1.0000
-0.250 0.1021 0.01381 0.00374 -0.0255 1.0000 1.0000
0.000 0.1278 0.01385 0.00364 -0.0250 1.0000 1.0000
0.250 0.1532 0.01391 0.00360 -0.0246 1.0000 1.0000
0.500 0.1785 0.01398 0.00361 -0.0241 1.0000 1.0000
0.750 0.2035 0.01407 0.00369 -0.0236 1.0000 1.0000
1.000 0.2282 0.01419 0.00381 -0.0232 1.0000 1.0000
1.250 0.2527 0.01433 0.00400 -0.0228 1.0000 1.0000
1.500 0.2770 0.01450 0.00426 -0.0224 1.0000 1.0000
1.750 0.3009 0.01473 0.00462 -0.0222 1.0000 1.0000
2.000 0.3242 0.01502 0.00507 -0.0220 1.0000 1.0000
2.250 0.3464 0.01546 0.00569 -0.0222 1.0000 1.0000
2.500 0.3663 0.01628 0.00673 -0.0233 1.0000 1.0000
2.750 0.4932 0.01628 0.00733 -0.0409 0.8670 1.0000
3.000 0.5206 0.01669 0.00747 -0.0372 0.7550 1.0000
3.250 0.5388 0.01747 0.00788 -0.0329 0.6699 1.0000
3.500 0.5594 0.01839 0.00842 -0.0298 0.5994 1.0000
3.750 0.5823 0.01939 0.00915 -0.0278 0.5369 1.0000
4.000 0.6057 0.02045 0.01000 -0.0261 0.4820 1.0000
4.250 0.6300 0.02161 0.01098 -0.0247 0.4320 1.0000
4.500 0.6545 0.02283 0.01214 -0.0236 0.3853 1.0000
4.750 0.6790 0.02412 0.01332 -0.0224 0.3424 1.0000
5.000 0.7034 0.02546 0.01456 -0.0213 0.3016 1.0000
5.250 0.7280 0.02697 0.01613 -0.0204 0.2619 1.0000
5.500 0.7520 0.02857 0.01760 -0.0193 0.2269 1.0000
5.750 0.7765 0.03056 0.01982 -0.0185 0.1945 1.0000
6.000 0.8005 0.03274 0.02200 -0.0176 0.1688 1.0000
6.250 0.8244 0.03573 0.02551 -0.0170 0.1492 1.0000
6.500 0.8469 0.03903 0.02925 -0.0165 0.1351 1.0000
6.750 0.8679 0.04263 0.03317 -0.0160 0.1251 1.0000
7.000 0.8854 0.04765 0.03894 -0.0162 0.1198 1.0000
7.250 0.9006 0.05286 0.04464 -0.0165 0.1174 1.0000
7.500 0.9150 0.05754 0.04949 -0.0165 0.1135 1.0000
7.750 0.9227 0.06390 0.05630 -0.0178 0.1137 1.0000
8.000 0.9097 0.07524 0.06835 -0.0247 0.1252 1.0000
8.250 0.9218 0.08068 0.07380 -0.0244 0.1277 1.0000
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Polar data table (+)
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