AG09 (ag09-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: AG09 (ag09-il) Reynolds number: 50,000 Max Cl/Cd: 32.35 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ag09-il-50000.txt Download as CSV file: xf-ag09-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: AG09
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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5995 0.11592 0.10931 0.0211 1.0000 0.1507
-8.000 -0.5788 0.10908 0.10244 0.0242 1.0000 0.1571
-7.750 -0.5805 0.10680 0.10024 0.0211 1.0000 0.1634
-7.500 -0.5757 0.10273 0.09624 0.0196 1.0000 0.1675
-7.250 -0.5680 0.09923 0.09276 0.0182 1.0000 0.1768
-7.000 -0.5595 0.09503 0.08857 0.0173 1.0000 0.1843
-6.750 -0.5557 0.09220 0.08580 0.0101 1.0000 0.1941
-6.500 -0.5464 0.08901 0.08261 0.0059 1.0000 0.2074
-6.000 -0.5236 0.08049 0.07415 0.0061 1.0000 0.2365
-5.750 -0.5112 0.07636 0.07006 0.0069 1.0000 0.2521
-5.500 -0.4984 0.07258 0.06631 0.0074 1.0000 0.2708
-5.250 -0.4863 0.06894 0.06268 0.0062 1.0000 0.2952
-4.750 -0.4637 0.06220 0.05602 0.0083 1.0000 0.3638
-4.500 -0.4560 0.05906 0.05296 0.0147 1.0000 0.4121
-4.000 -0.4488 0.05356 0.04769 0.0340 1.0000 0.5547
-3.750 -0.2897 0.03954 0.03127 -0.0310 1.0000 0.1765
-3.500 -0.2514 0.03467 0.02560 -0.0337 1.0000 0.1479
-3.250 -0.2198 0.03152 0.02184 -0.0346 1.0000 0.1496
-3.000 -0.1881 0.02864 0.01829 -0.0349 1.0000 0.1510
-2.750 -0.1604 0.02604 0.01556 -0.0348 1.0000 0.1586
-2.500 -0.1316 0.02409 0.01322 -0.0345 1.0000 0.1761
-2.250 -0.1023 0.02221 0.01102 -0.0341 1.0000 0.1937
-2.000 -0.0755 0.02061 0.00937 -0.0334 1.0000 0.2225
-1.750 -0.0473 0.01916 0.00792 -0.0327 1.0000 0.2637
-1.500 -0.0199 0.01757 0.00661 -0.0320 1.0000 0.3230
-1.250 0.0074 0.01374 0.00510 -0.0299 1.0000 1.0000
-1.000 0.0349 0.01375 0.00440 -0.0293 1.0000 1.0000
-0.750 0.0606 0.01376 0.00406 -0.0288 1.0000 1.0000
-0.500 0.0859 0.01379 0.00384 -0.0283 1.0000 1.0000
-0.250 0.1109 0.01384 0.00370 -0.0278 1.0000 1.0000
0.000 0.1356 0.01390 0.00361 -0.0272 1.0000 1.0000
0.250 0.1600 0.01398 0.00359 -0.0267 1.0000 1.0000
0.500 0.1843 0.01409 0.00364 -0.0262 1.0000 1.0000
0.750 0.2083 0.01422 0.00375 -0.0258 1.0000 1.0000
1.000 0.2320 0.01439 0.00392 -0.0254 1.0000 1.0000
1.250 0.2553 0.01460 0.00416 -0.0251 1.0000 1.0000
1.500 0.2782 0.01487 0.00448 -0.0248 1.0000 1.0000
1.750 0.3005 0.01522 0.00492 -0.0247 1.0000 1.0000
2.000 0.3223 0.01569 0.00549 -0.0249 1.0000 1.0000
2.250 0.3427 0.01636 0.00630 -0.0256 1.0000 1.0000
2.500 0.3818 0.01730 0.00745 -0.0306 0.9866 1.0000
2.750 0.4965 0.01737 0.00811 -0.0463 0.9076 1.0000
3.000 0.5450 0.01727 0.00816 -0.0469 0.8295 1.0000
3.250 0.5668 0.01757 0.00833 -0.0426 0.7574 1.0000
3.500 0.5856 0.01810 0.00865 -0.0385 0.6894 1.0000
3.750 0.6059 0.01879 0.00918 -0.0353 0.6253 1.0000
4.000 0.6275 0.01959 0.00978 -0.0327 0.5672 1.0000
4.250 0.6502 0.02050 0.01051 -0.0307 0.5143 1.0000
4.500 0.6737 0.02151 0.01137 -0.0290 0.4656 1.0000
4.750 0.6973 0.02260 0.01230 -0.0274 0.4206 1.0000
5.000 0.7210 0.02379 0.01352 -0.0261 0.3756 1.0000
5.250 0.7444 0.02506 0.01475 -0.0248 0.3315 1.0000
5.500 0.7674 0.02643 0.01607 -0.0235 0.2868 1.0000
5.750 0.7899 0.02791 0.01743 -0.0221 0.2416 1.0000
6.000 0.8121 0.02963 0.01900 -0.0208 0.1984 1.0000
6.250 0.8347 0.03193 0.02155 -0.0197 0.1627 1.0000
6.500 0.8577 0.03482 0.02460 -0.0187 0.1400 1.0000
6.750 0.8795 0.03790 0.02810 -0.0178 0.1226 1.0000
7.000 0.9004 0.04183 0.03237 -0.0171 0.1140 1.0000
7.250 0.9200 0.04505 0.03583 -0.0164 0.1039 1.0000
7.500 0.9340 0.05046 0.04192 -0.0162 0.1017 1.0000
7.750 0.9444 0.05640 0.04846 -0.0164 0.1018 1.0000
8.000 0.9534 0.06243 0.05486 -0.0168 0.1027 1.0000
8.250 0.9335 0.07294 0.06628 -0.0225 0.1113 1.0000
8.500 0.9348 0.07937 0.07281 -0.0242 0.1147 1.0000
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Polar data table (+)
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