OAF095 AIRFOIL (oaf095-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: OAF095 AIRFOIL (oaf095-il) Reynolds number: 100,000 Max Cl/Cd: 63.95 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-oaf095-il-100000-n5.txt Download as CSV file: xf-oaf095-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: OAF095 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4703 0.08778 0.08257 -0.0347 1.0000 0.0407
-9.250 -0.4629 0.08543 0.08025 -0.0358 1.0000 0.0420
-9.000 -0.5688 0.04633 0.04087 -0.0785 1.0000 0.0392
-8.750 -0.5567 0.04072 0.03493 -0.0860 1.0000 0.0405
-8.500 -0.5390 0.03859 0.03271 -0.0879 1.0000 0.0421
-8.250 -0.5187 0.03506 0.02879 -0.0918 1.0000 0.0441
-8.000 -0.4951 0.03143 0.02460 -0.0954 1.0000 0.0461
-7.750 -0.4704 0.02880 0.02153 -0.0976 1.0000 0.0487
-7.500 -0.4486 0.02777 0.02049 -0.0979 1.0000 0.0512
-7.250 -0.4245 0.02634 0.01884 -0.0987 1.0000 0.0543
-7.000 -0.3987 0.02476 0.01685 -0.0996 1.0000 0.0576
-6.750 -0.3681 0.02373 0.01582 -0.1013 0.9978 0.0617
-6.500 -0.3300 0.02254 0.01438 -0.1042 0.9941 0.0673
-6.250 -0.2933 0.02154 0.01329 -0.1069 0.9897 0.0733
-6.000 -0.2552 0.02064 0.01215 -0.1095 0.9857 0.0814
-5.750 -0.2171 0.01987 0.01136 -0.1122 0.9821 0.0906
-5.500 -0.1805 0.01919 0.01060 -0.1145 0.9771 0.1012
-5.250 -0.1421 0.01860 0.00995 -0.1171 0.9731 0.1144
-5.000 -0.1029 0.01811 0.00940 -0.1198 0.9696 0.1295
-4.750 -0.0678 0.01773 0.00897 -0.1216 0.9636 0.1453
-4.500 -0.0303 0.01738 0.00855 -0.1237 0.9589 0.1629
-4.250 0.0057 0.01708 0.00819 -0.1256 0.9536 0.1804
-4.000 0.0397 0.01683 0.00792 -0.1270 0.9470 0.1973
-3.750 0.0762 0.01657 0.00767 -0.1288 0.9422 0.2148
-3.500 0.1077 0.01640 0.00749 -0.1296 0.9345 0.2310
-3.250 0.1418 0.01621 0.00728 -0.1308 0.9284 0.2481
-3.000 0.1731 0.01607 0.00714 -0.1314 0.9209 0.2645
-2.750 0.2051 0.01593 0.00701 -0.1321 0.9138 0.2816
-2.500 0.2358 0.01582 0.00690 -0.1326 0.9063 0.2990
-2.250 0.2661 0.01572 0.00681 -0.1329 0.8985 0.3168
-2.000 0.2960 0.01565 0.00674 -0.1331 0.8908 0.3354
-1.750 0.3254 0.01557 0.00668 -0.1332 0.8826 0.3537
-1.500 0.3543 0.01551 0.00666 -0.1332 0.8746 0.3716
-1.250 0.3833 0.01545 0.00663 -0.1331 0.8667 0.3895
-1.000 0.4118 0.01541 0.00663 -0.1330 0.8585 0.4076
-0.750 0.4404 0.01535 0.00661 -0.1329 0.8507 0.4256
-0.500 0.4685 0.01533 0.00664 -0.1327 0.8424 0.4440
0.000 0.5248 0.01527 0.00670 -0.1323 0.8264 0.4831
0.250 0.5529 0.01523 0.00673 -0.1320 0.8187 0.5049
0.500 0.5805 0.01523 0.00682 -0.1317 0.8098 0.5286
0.750 0.6084 0.01516 0.00685 -0.1313 0.8025 0.5558
1.000 0.6356 0.01515 0.00698 -0.1310 0.7930 0.5864
1.250 0.6628 0.01502 0.00701 -0.1303 0.7853 0.6240
1.500 0.6883 0.01492 0.00715 -0.1295 0.7747 0.6766
1.750 0.7091 0.01461 0.00719 -0.1273 0.7649 0.7800
2.000 0.7312 0.01436 0.00709 -0.1254 0.7556 1.0000
2.250 0.7595 0.01453 0.00724 -0.1253 0.7443 1.0000
2.500 0.7878 0.01465 0.00735 -0.1251 0.7331 1.0000
2.750 0.8156 0.01474 0.00743 -0.1246 0.7199 1.0000
3.000 0.8429 0.01482 0.00749 -0.1240 0.7040 1.0000
3.250 0.8700 0.01491 0.00757 -0.1234 0.6863 1.0000
3.500 0.8969 0.01501 0.00766 -0.1227 0.6673 1.0000
3.750 0.9236 0.01511 0.00773 -0.1219 0.6466 1.0000
4.000 0.9496 0.01527 0.00788 -0.1211 0.6208 1.0000
4.250 0.9754 0.01545 0.00802 -0.1202 0.5918 1.0000
4.500 1.0006 0.01570 0.00820 -0.1193 0.5576 1.0000
4.750 1.0251 0.01603 0.00842 -0.1183 0.5154 1.0000
5.000 1.0482 0.01652 0.00869 -0.1171 0.4606 1.0000
5.250 1.0689 0.01732 0.00912 -0.1158 0.3808 1.0000
5.500 1.0864 0.01858 0.00982 -0.1143 0.2913 1.0000
5.750 1.1037 0.01997 0.01073 -0.1130 0.2166 1.0000
6.000 1.1221 0.02125 0.01169 -0.1118 0.1643 1.0000
6.250 1.1408 0.02248 0.01269 -0.1106 0.1293 1.0000
6.500 1.1600 0.02359 0.01370 -0.1094 0.1066 1.0000
6.750 1.1787 0.02472 0.01476 -0.1082 0.0915 1.0000
7.000 1.1972 0.02583 0.01587 -0.1068 0.0810 1.0000
7.250 1.2145 0.02703 0.01705 -0.1054 0.0731 1.0000
7.500 1.2321 0.02816 0.01825 -0.1039 0.0666 1.0000
7.750 1.2477 0.02948 0.01959 -0.1023 0.0620 1.0000
8.000 1.2647 0.03066 0.02087 -0.1007 0.0573 1.0000
8.250 1.2795 0.03201 0.02221 -0.0991 0.0537 1.0000
8.750 1.3116 0.03483 0.02528 -0.0958 0.0481 1.0000
9.000 1.3266 0.03620 0.02670 -0.0942 0.0454 1.0000
9.250 1.3403 0.03791 0.02835 -0.0926 0.0433 1.0000
9.500 1.3573 0.03952 0.03021 -0.0911 0.0415 1.0000
9.750 1.3735 0.04131 0.03219 -0.0896 0.0398 1.0000
10.000 1.3875 0.04304 0.03410 -0.0880 0.0382 1.0000
10.250 1.3981 0.04466 0.03581 -0.0860 0.0367 1.0000
10.500 1.4088 0.04655 0.03771 -0.0843 0.0354 1.0000
10.750 1.4180 0.04882 0.04026 -0.0823 0.0343 1.0000
11.000 1.4252 0.05136 0.04313 -0.0802 0.0334 1.0000
11.250 1.4292 0.05408 0.04615 -0.0780 0.0326 1.0000
11.500 1.4298 0.05691 0.04930 -0.0758 0.0319 1.0000
11.750 1.4276 0.05981 0.05246 -0.0738 0.0312 1.0000
12.000 1.4235 0.06279 0.05568 -0.0720 0.0305 1.0000
12.250 1.4183 0.06587 0.05897 -0.0705 0.0300 1.0000
12.500 1.4124 0.06908 0.06237 -0.0694 0.0294 1.0000
12.750 1.4059 0.07252 0.06596 -0.0686 0.0290 1.0000
13.000 1.3990 0.07620 0.06977 -0.0682 0.0285 1.0000
13.250 1.3853 0.08096 0.07473 -0.0682 0.0283 1.0000
13.500 1.3643 0.08682 0.08092 -0.0693 0.0282 1.0000
13.750 1.3419 0.09334 0.08773 -0.0714 0.0281 1.0000
14.000 1.3188 0.10050 0.09516 -0.0744 0.0281 1.0000
14.250 1.2943 0.10851 0.10342 -0.0786 0.0281 1.0000
14.500 1.2693 0.11752 0.11264 -0.0840 0.0282 1.0000
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Polar data table (+)
Polar graphs
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