OAF102 AIRFOIL (oaf102-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: OAF102 AIRFOIL (oaf102-il) Reynolds number: 500,000 Max Cl/Cd: 86.33 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-oaf102-il-500000-n5.txt Download as CSV file: xf-oaf102-il-500000-n5.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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.750 -0.7989 0.10012 0.09750 -0.0113 1.0000 0.0097
-13.500 -0.8319 0.08909 0.08634 -0.0177 1.0000 0.0096
-13.250 -0.8579 0.07995 0.07708 -0.0234 1.0000 0.0096
-13.000 -0.8820 0.07163 0.06865 -0.0290 1.0000 0.0095
-12.750 -0.9041 0.06379 0.06070 -0.0346 1.0000 0.0095
-12.500 -0.9255 0.05629 0.05308 -0.0402 1.0000 0.0094
-12.250 -0.9423 0.04840 0.04507 -0.0481 1.0000 0.0094
-12.000 -0.9405 0.03921 0.03565 -0.0619 1.0000 0.0095
-11.750 -0.9161 0.03133 0.02745 -0.0768 1.0000 0.0097
-11.500 -0.8904 0.02838 0.02430 -0.0819 1.0000 0.0100
-11.250 -0.8639 0.02635 0.02209 -0.0852 1.0000 0.0102
-11.000 -0.8365 0.02468 0.02027 -0.0877 1.0000 0.0106
-10.750 -0.8085 0.02322 0.01865 -0.0899 1.0000 0.0110
-10.500 -0.7800 0.02192 0.01718 -0.0918 1.0000 0.0114
-10.250 -0.7509 0.02075 0.01586 -0.0935 1.0000 0.0118
-10.000 -0.7216 0.01975 0.01472 -0.0950 1.0000 0.0122
-9.750 -0.6916 0.01855 0.01341 -0.0969 1.0000 0.0128
-9.500 -0.6604 0.01772 0.01249 -0.0985 0.9877 0.0135
-9.250 -0.6279 0.01705 0.01172 -0.1002 0.9735 0.0143
-9.000 -0.5972 0.01645 0.01101 -0.1014 0.9613 0.0152
-8.750 -0.5681 0.01589 0.01032 -0.1021 0.9500 0.0160
-8.500 -0.5396 0.01521 0.00955 -0.1028 0.9392 0.0171
-8.250 -0.5115 0.01474 0.00899 -0.1031 0.9284 0.0182
-8.000 -0.4830 0.01430 0.00845 -0.1035 0.9183 0.0194
-7.750 -0.4545 0.01389 0.00793 -0.1039 0.9089 0.0206
-7.500 -0.4255 0.01342 0.00740 -0.1044 0.8990 0.0226
-7.250 -0.3965 0.01308 0.00697 -0.1048 0.8893 0.0246
-7.000 -0.3674 0.01274 0.00652 -0.1052 0.8801 0.0265
-6.750 -0.3378 0.01236 0.00610 -0.1058 0.8706 0.0295
-6.500 -0.3084 0.01209 0.00574 -0.1061 0.8614 0.0324
-6.250 -0.2788 0.01176 0.00535 -0.1067 0.8519 0.0365
-6.000 -0.2490 0.01151 0.00503 -0.1071 0.8426 0.0407
-5.750 -0.2192 0.01122 0.00470 -0.1076 0.8340 0.0473
-5.500 -0.1892 0.01097 0.00441 -0.1081 0.8247 0.0549
-5.250 -0.1593 0.01072 0.00414 -0.1087 0.8162 0.0651
-5.000 -0.1293 0.01048 0.00389 -0.1092 0.8075 0.0785
-4.750 -0.0992 0.01025 0.00367 -0.1097 0.7994 0.0936
-4.500 -0.0692 0.01005 0.00347 -0.1103 0.7912 0.1109
-4.250 -0.0392 0.00987 0.00329 -0.1108 0.7833 0.1281
-4.000 -0.0093 0.00971 0.00313 -0.1112 0.7753 0.1468
-3.750 0.0207 0.00956 0.00301 -0.1117 0.7666 0.1671
-3.500 0.0504 0.00947 0.00290 -0.1121 0.7585 0.1839
-3.250 0.0803 0.00937 0.00279 -0.1125 0.7503 0.1985
-3.000 0.1101 0.00929 0.00270 -0.1129 0.7437 0.2132
-2.750 0.1400 0.00920 0.00263 -0.1133 0.7368 0.2281
-2.500 0.1698 0.00913 0.00256 -0.1137 0.7305 0.2427
-2.250 0.1997 0.00907 0.00251 -0.1141 0.7248 0.2571
-2.000 0.2295 0.00901 0.00247 -0.1144 0.7188 0.2711
-1.750 0.2592 0.00898 0.00243 -0.1148 0.7134 0.2853
-1.500 0.2890 0.00893 0.00241 -0.1151 0.7073 0.2994
-1.250 0.3186 0.00891 0.00240 -0.1154 0.7013 0.3136
-1.000 0.3483 0.00889 0.00239 -0.1158 0.6959 0.3276
-0.750 0.3780 0.00887 0.00240 -0.1161 0.6903 0.3411
-0.500 0.4075 0.00888 0.00241 -0.1164 0.6847 0.3549
-0.250 0.4372 0.00887 0.00244 -0.1167 0.6783 0.3693
0.000 0.4667 0.00887 0.00246 -0.1169 0.6717 0.3826
0.250 0.4962 0.00888 0.00250 -0.1172 0.6660 0.3937
0.500 0.5258 0.00889 0.00254 -0.1175 0.6599 0.4038
1.000 0.5847 0.00893 0.00264 -0.1180 0.6482 0.4249
1.250 0.6141 0.00896 0.00270 -0.1183 0.6420 0.4354
1.500 0.6435 0.00899 0.00277 -0.1185 0.6361 0.4472
1.750 0.6730 0.00902 0.00285 -0.1188 0.6296 0.4598
2.250 0.7317 0.00908 0.00304 -0.1193 0.6164 0.4849
2.500 0.7607 0.00914 0.00312 -0.1195 0.6027 0.5009
2.750 0.7893 0.00925 0.00319 -0.1196 0.5706 0.5205
3.000 0.8175 0.00947 0.00329 -0.1197 0.5237 0.5414
3.250 0.8454 0.00980 0.00347 -0.1199 0.4737 0.5640
3.500 0.8734 0.01014 0.00370 -0.1201 0.4280 0.5895
3.750 0.9004 0.01074 0.00406 -0.1203 0.3504 0.6235
4.000 0.9256 0.01174 0.00464 -0.1205 0.2366 0.6756
4.250 0.9461 0.01226 0.00523 -0.1194 0.1480 0.8674
4.500 0.9674 0.01274 0.00554 -0.1182 0.0979 1.0000
4.750 0.9940 0.01329 0.00594 -0.1182 0.0703 1.0000
5.000 1.0208 0.01375 0.00631 -0.1182 0.0550 1.0000
5.250 1.0476 0.01417 0.00668 -0.1181 0.0456 1.0000
5.500 1.0745 0.01455 0.00706 -0.1180 0.0396 1.0000
5.750 1.1009 0.01498 0.00746 -0.1179 0.0344 1.0000
6.000 1.1272 0.01539 0.00787 -0.1177 0.0307 1.0000
6.250 1.1530 0.01587 0.00834 -0.1175 0.0277 1.0000
6.500 1.1789 0.01628 0.00880 -0.1173 0.0256 1.0000
6.750 1.2042 0.01676 0.00927 -0.1170 0.0234 1.0000
7.000 1.2288 0.01732 0.00984 -0.1165 0.0215 1.0000
7.250 1.2536 0.01781 0.01037 -0.1161 0.0203 1.0000
7.500 1.2778 0.01833 0.01094 -0.1156 0.0191 1.0000
7.750 1.3013 0.01891 0.01154 -0.1151 0.0180 1.0000
8.000 1.3232 0.01967 0.01232 -0.1142 0.0169 1.0000
8.250 1.3462 0.02023 0.01294 -0.1136 0.0161 1.0000
8.500 1.3684 0.02085 0.01363 -0.1128 0.0153 1.0000
8.750 1.3898 0.02151 0.01434 -0.1119 0.0146 1.0000
9.000 1.4104 0.02221 0.01508 -0.1109 0.0140 1.0000
9.250 1.4295 0.02302 0.01593 -0.1097 0.0135 1.0000
9.500 1.4453 0.02408 0.01705 -0.1080 0.0129 1.0000
9.750 1.4626 0.02490 0.01796 -0.1066 0.0126 1.0000
10.000 1.4783 0.02580 0.01896 -0.1048 0.0122 1.0000
10.250 1.4927 0.02673 0.01998 -0.1030 0.0119 1.0000
10.500 1.5059 0.02768 0.02102 -0.1010 0.0115 1.0000
10.750 1.5165 0.02864 0.02205 -0.0986 0.0111 1.0000
11.000 1.5259 0.02971 0.02319 -0.0962 0.0108 1.0000
11.250 1.5354 0.03088 0.02444 -0.0941 0.0105 1.0000
11.500 1.5433 0.03227 0.02591 -0.0920 0.0102 1.0000
11.750 1.5491 0.03395 0.02769 -0.0899 0.0100 1.0000
12.000 1.5525 0.03594 0.02978 -0.0878 0.0097 1.0000
12.250 1.5596 0.03766 0.03163 -0.0862 0.0096 1.0000
12.500 1.5657 0.03953 0.03363 -0.0846 0.0094 1.0000
12.750 1.5708 0.04154 0.03578 -0.0832 0.0093 1.0000
13.000 1.5750 0.04372 0.03810 -0.0818 0.0091 1.0000
13.250 1.5779 0.04606 0.04058 -0.0805 0.0089 1.0000
13.500 1.5801 0.04856 0.04322 -0.0794 0.0088 1.0000
13.750 1.5812 0.05123 0.04603 -0.0785 0.0086 1.0000
14.000 1.5812 0.05407 0.04901 -0.0776 0.0085 1.0000
14.250 1.5801 0.05712 0.05221 -0.0770 0.0083 1.0000
14.500 1.5783 0.06033 0.05557 -0.0766 0.0082 1.0000
14.750 1.5754 0.06374 0.05911 -0.0763 0.0081 1.0000
15.000 1.5715 0.06736 0.06286 -0.0763 0.0080 1.0000
15.250 1.5668 0.07124 0.06687 -0.0765 0.0079 1.0000
15.500 1.5615 0.07537 0.07114 -0.0771 0.0078 1.0000
15.750 1.5546 0.07987 0.07577 -0.0779 0.0077 1.0000
16.000 1.5465 0.08474 0.08077 -0.0791 0.0076 1.0000
16.250 1.5367 0.09008 0.08626 -0.0807 0.0075 1.0000
16.500 1.5255 0.09587 0.09219 -0.0827 0.0075 1.0000
16.750 1.5125 0.10216 0.09863 -0.0851 0.0074 1.0000
17.000 1.4974 0.10908 0.10571 -0.0881 0.0074 1.0000
17.250 1.4808 0.11660 0.11338 -0.0917 0.0073 1.0000
17.500 1.4630 0.12470 0.12165 -0.0959 0.0073 1.0000
17.750 1.4431 0.13361 0.13074 -0.1008 0.0073 1.0000
18.000 1.4218 0.14335 0.14066 -0.1066 0.0073 1.0000
18.250 1.3973 0.15457 0.15208 -0.1137 0.0073 1.0000
18.500 1.3609 0.17014 0.16791 -0.1238 0.0074 1.0000
|
Polar data table (+)
Polar graphs
<< Back to OAF102 AIRFOIL (oaf102-il)