OAF102 AIRFOIL (oaf102-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: OAF102 AIRFOIL (oaf102-il) Reynolds number: 50,000 Max Cl/Cd: 40.68 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oaf102-il-50000.txt Download as CSV file: xf-oaf102-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: OAF102 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4243 0.10953 0.10257 -0.0158 1.0000 0.2608
-8.250 -0.3968 0.10352 0.09648 -0.0146 1.0000 0.2706
-8.000 -0.4043 0.10141 0.09449 -0.0163 1.0000 0.2776
-7.750 -0.3954 0.09825 0.09134 -0.0163 1.0000 0.2895
-7.500 -0.3863 0.09424 0.08736 -0.0165 1.0000 0.2952
-7.250 -0.3887 0.09205 0.08526 -0.0169 1.0000 0.3045
-7.000 -0.3802 0.08829 0.08154 -0.0167 1.0000 0.3114
-6.750 -0.4278 0.06840 0.06179 -0.0540 1.0000 0.1714
-6.500 -0.4189 0.06741 0.06088 -0.0482 1.0000 0.1748
-6.250 -0.4091 0.05757 0.05086 -0.0603 1.0000 0.1671
-6.000 -0.3787 0.04581 0.03840 -0.0767 1.0000 0.1695
-5.750 -0.3557 0.04196 0.03433 -0.0795 1.0000 0.1768
-5.500 -0.3286 0.03847 0.03046 -0.0830 1.0000 0.1910
-5.250 -0.3013 0.03567 0.02735 -0.0853 1.0000 0.2066
-5.000 -0.2745 0.03377 0.02516 -0.0870 1.0000 0.2287
-4.750 -0.2530 0.03284 0.02421 -0.0866 1.0000 0.2499
-4.500 -0.2275 0.03178 0.02294 -0.0872 1.0000 0.2772
-4.250 -0.2113 0.03189 0.02318 -0.0850 1.0000 0.2994
-4.000 -0.1892 0.03145 0.02267 -0.0845 1.0000 0.3260
-3.750 -0.1656 0.03101 0.02210 -0.0845 1.0000 0.3538
-3.500 -0.1467 0.03094 0.02206 -0.0831 1.0000 0.3767
-3.250 -0.1232 0.03059 0.02160 -0.0831 1.0000 0.4023
-3.000 -0.0962 0.03017 0.02100 -0.0840 1.0000 0.4294
-2.750 -0.0758 0.03001 0.02087 -0.0831 1.0000 0.4498
-2.500 -0.0487 0.02969 0.02037 -0.0842 1.0000 0.4755
-2.250 -0.0269 0.02951 0.02020 -0.0837 1.0000 0.4957
-2.000 0.0003 0.02932 0.01987 -0.0849 1.0000 0.5208
-1.750 0.0226 0.02920 0.01978 -0.0847 1.0000 0.5417
-1.500 0.0494 0.02912 0.01960 -0.0858 1.0000 0.5668
-1.250 0.0721 0.02910 0.01960 -0.0858 1.0000 0.5895
-1.000 0.0974 0.02913 0.01960 -0.0866 1.0000 0.6150
-0.750 0.1227 0.02924 0.01967 -0.0873 1.0000 0.6414
-0.500 0.1460 0.02938 0.01982 -0.0876 1.0000 0.6672
-0.250 0.1697 0.02957 0.02005 -0.0880 1.0000 0.6957
0.000 0.1928 0.02981 0.02034 -0.0883 1.0000 0.7274
0.250 0.2151 0.03007 0.02070 -0.0884 1.0000 0.7647
0.500 0.2396 0.03015 0.02098 -0.0885 0.9967 0.8139
0.750 0.2583 0.02973 0.02088 -0.0875 0.9874 0.9371
1.000 0.3169 0.03074 0.02160 -0.0973 0.9772 1.0000
1.250 0.3712 0.03186 0.02244 -0.1048 0.9683 1.0000
1.500 0.4135 0.03290 0.02331 -0.1094 0.9587 1.0000
1.750 0.4480 0.03396 0.02427 -0.1124 0.9488 1.0000
2.000 0.4892 0.03506 0.02530 -0.1162 0.9396 1.0000
2.250 0.5219 0.03615 0.02637 -0.1186 0.9296 1.0000
2.500 0.5504 0.03730 0.02753 -0.1202 0.9193 1.0000
2.750 0.5878 0.03846 0.02871 -0.1231 0.9098 1.0000
3.000 0.6163 0.03963 0.02994 -0.1245 0.8991 1.0000
3.250 0.6402 0.04091 0.03130 -0.1252 0.8882 1.0000
3.500 0.6720 0.04215 0.03263 -0.1270 0.8777 1.0000
3.750 0.7077 0.04331 0.03391 -0.1293 0.8667 1.0000
4.250 0.7509 0.04606 0.03693 -0.1296 0.8416 1.0000
4.500 0.7779 0.04736 0.03838 -0.1304 0.8282 1.0000
4.750 0.8061 0.04857 0.03977 -0.1310 0.8137 1.0000
5.000 0.8354 0.04963 0.04108 -0.1315 0.7974 1.0000
5.250 0.8846 0.04949 0.04125 -0.1331 0.7763 1.0000
5.500 1.0131 0.03792 0.03035 -0.1292 0.7116 1.0000
5.750 1.0585 0.03295 0.02571 -0.1227 0.6721 1.0000
6.000 1.0907 0.02779 0.02078 -0.1137 0.6075 1.0000
6.250 1.0829 0.02662 0.01796 -0.1012 0.3239 1.0000
6.500 1.0763 0.03066 0.02051 -0.0966 0.2229 1.0000
6.750 1.0884 0.03331 0.02273 -0.0940 0.1849 1.0000
7.000 1.1153 0.03560 0.02472 -0.0928 0.1613 1.0000
7.250 1.1491 0.03785 0.02688 -0.0924 0.1433 1.0000
7.500 1.1867 0.04061 0.02956 -0.0927 0.1327 1.0000
7.750 1.2178 0.04340 0.03236 -0.0926 0.1236 1.0000
8.000 1.2440 0.04657 0.03592 -0.0918 0.1187 1.0000
8.250 1.2678 0.04999 0.03985 -0.0906 0.1162 1.0000
8.500 1.2876 0.05349 0.04375 -0.0894 0.1133 1.0000
8.750 1.3093 0.05733 0.04767 -0.0887 0.1095 1.0000
9.000 1.3251 0.06197 0.05266 -0.0875 0.1086 1.0000
9.250 1.3244 0.06588 0.05746 -0.0843 0.1106 1.0000
9.500 1.3179 0.07095 0.06325 -0.0813 0.1135 1.0000
9.750 1.3081 0.07622 0.06903 -0.0788 0.1163 1.0000
10.000 1.2986 0.08150 0.07465 -0.0768 0.1186 1.0000
10.250 1.2931 0.08703 0.08039 -0.0753 0.1205 1.0000
10.500 1.2706 0.09195 0.08565 -0.0733 0.1237 1.0000
10.750 1.2022 0.09779 0.09183 -0.0722 0.1268 1.0000
11.000 1.1540 0.10648 0.10067 -0.0760 0.1297 1.0000
11.250 1.1375 0.11464 0.10886 -0.0790 0.1342 1.0000
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Polar data table (+)
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