20-32C AIRFOIL (2032c-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: 20-32C AIRFOIL (2032c-il) Reynolds number: 50,000 Max Cl/Cd: 31.7 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-2032c-il-50000.txt Download as CSV file: xf-2032c-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: 20-32C AIRFOIL
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 -0.2446 0.11407 0.10788 -0.0201 1.0000 0.0956
-8.000 -0.2475 0.11417 0.10812 -0.0218 1.0000 0.0969
-7.750 -0.2530 0.11491 0.10902 -0.0240 1.0000 0.0975
-7.500 -0.2326 0.10640 0.10052 -0.0214 1.0000 0.1007
-7.250 -0.2270 0.10345 0.09765 -0.0208 1.0000 0.1042
-7.000 -0.2270 0.10172 0.09604 -0.0205 1.0000 0.1074
-6.750 -0.2298 0.10096 0.09544 -0.0212 1.0000 0.1103
-6.500 -0.2352 0.10161 0.09625 -0.0237 1.0000 0.1117
-6.250 -0.2376 0.10049 0.09527 -0.0248 1.0000 0.1126
-6.000 -0.2347 0.09499 0.08985 -0.0189 1.0000 0.1159
-5.750 -0.2378 0.09320 0.08817 -0.0171 1.0000 0.1190
-5.500 -0.2425 0.09209 0.08716 -0.0165 1.0000 0.1224
-5.250 -0.2448 0.09223 0.08740 -0.0194 1.0000 0.1257
-4.750 -0.2495 0.08760 0.08296 -0.0165 1.0000 0.1307
-4.500 -0.2491 0.08593 0.08136 -0.0163 1.0000 0.1361
-4.250 -0.2347 0.08537 0.08078 -0.0238 1.0000 0.1420
-4.000 -0.2399 0.08224 0.07777 -0.0193 1.0000 0.1460
-3.750 -0.2119 0.08076 0.07624 -0.0284 0.9978 0.1567
-3.500 -0.1606 0.07544 0.07086 -0.0374 0.9845 0.1722
-3.250 -0.1097 0.07057 0.06593 -0.0457 0.9703 0.1906
-3.000 -0.0468 0.06613 0.06135 -0.0577 0.9549 0.2171
-2.750 0.0020 0.06206 0.05719 -0.0647 0.9387 0.2506
-2.500 0.0517 0.05810 0.05317 -0.0716 0.9220 0.2936
-2.250 0.0979 0.05414 0.04917 -0.0769 0.9059 0.3480
-1.500 0.1619 0.04319 0.03847 -0.0674 0.8612 0.6068
-1.000 0.4963 0.03560 0.02784 -0.1339 0.8226 0.1971
-0.750 0.5400 0.03345 0.02515 -0.1356 0.8029 0.1801
-0.500 0.5796 0.03183 0.02304 -0.1363 0.7844 0.1823
-0.250 0.6150 0.03048 0.02121 -0.1360 0.7658 0.1822
0.000 0.6454 0.02963 0.01993 -0.1352 0.7447 0.1842
0.250 0.6759 0.02875 0.01871 -0.1339 0.7262 0.1919
0.500 0.7044 0.02821 0.01787 -0.1325 0.7078 0.2114
0.750 0.7304 0.02789 0.01743 -0.1312 0.6880 0.2430
1.000 0.7575 0.02681 0.01666 -0.1299 0.6710 0.3477
1.250 0.7812 0.02549 0.01587 -0.1275 0.6554 1.0000
1.500 0.8074 0.02596 0.01592 -0.1264 0.6392 1.0000
1.750 0.8333 0.02651 0.01613 -0.1255 0.6238 1.0000
2.000 0.8588 0.02713 0.01649 -0.1246 0.6090 1.0000
2.250 0.8838 0.02788 0.01702 -0.1240 0.5950 1.0000
2.500 0.9084 0.02871 0.01769 -0.1234 0.5815 1.0000
2.750 0.9327 0.02963 0.01849 -0.1228 0.5688 1.0000
3.000 0.9566 0.03064 0.01942 -0.1224 0.5572 1.0000
3.250 0.9819 0.03148 0.02013 -0.1218 0.5475 1.0000
3.500 1.0062 0.03245 0.02103 -0.1213 0.5374 1.0000
3.750 1.0274 0.03392 0.02253 -0.1211 0.5270 1.0000
4.000 1.0523 0.03485 0.02342 -0.1205 0.5190 1.0000
4.250 1.0728 0.03648 0.02512 -0.1203 0.5102 1.0000
4.500 1.0934 0.03811 0.02681 -0.1200 0.5025 1.0000
4.750 1.1149 0.03954 0.02826 -0.1196 0.4948 1.0000
5.000 1.1311 0.04178 0.03063 -0.1193 0.4873 1.0000
5.250 1.1500 0.04365 0.03264 -0.1190 0.4808 1.0000
5.500 1.1650 0.04615 0.03526 -0.1188 0.4750 1.0000
5.750 1.1687 0.05007 0.03941 -0.1188 0.4682 1.0000
6.000 1.2002 0.05025 0.03955 -0.1179 0.4635 1.0000
6.250 1.1521 0.06111 0.05079 -0.1194 0.4570 1.0000
6.500 1.0670 0.07694 0.06674 -0.1241 0.4575 1.0000
6.750 1.0401 0.08482 0.07461 -0.1262 0.4593 1.0000
7.000 1.0310 0.09057 0.08040 -0.1274 0.4611 1.0000
7.250 1.0388 0.09486 0.08479 -0.1282 0.4631 1.0000
7.500 0.9081 0.11777 0.10789 -0.1434 0.5936 1.0000
7.750 0.8981 0.11965 0.10978 -0.1415 0.5823 1.0000
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