DAE-11 AIRFOIL (dae11-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: DAE-11 AIRFOIL (dae11-il) Reynolds number: 200,000 Max Cl/Cd: 83.44 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-dae11-il-200000-n5.txt Download as CSV file: xf-dae11-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DAE-11 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.0500 0.09315 0.08837 -0.0736 0.7134 0.0288
-8.500 -0.0498 0.09040 0.08564 -0.0765 0.7116 0.0312
-8.250 -0.0471 0.08746 0.08271 -0.0788 0.7099 0.0314
-8.000 -0.0471 0.08427 0.07953 -0.0817 0.7083 0.0316
-7.500 -0.0401 0.07812 0.07347 -0.0851 0.7042 0.0317
-7.250 -0.0301 0.07542 0.07079 -0.0848 0.7020 0.0320
-7.000 -0.0226 0.07306 0.06845 -0.0852 0.6999 0.0323
-6.750 -0.0151 0.07021 0.06562 -0.0873 0.6979 0.0324
-6.500 -0.0116 0.06417 0.05957 -0.0936 0.6959 0.0266
-6.000 0.0114 0.05398 0.04923 -0.1035 0.6924 0.0217
-5.750 0.0287 0.04994 0.04509 -0.1077 0.6908 0.0214
-5.500 0.0484 0.04409 0.03909 -0.1136 0.6885 0.0214
-5.250 0.0693 0.03439 0.02873 -0.1206 0.6863 0.0218
-5.000 0.0925 0.03016 0.02400 -0.1228 0.6842 0.0218
-4.750 0.1174 0.02682 0.02010 -0.1241 0.6820 0.0218
-4.500 0.1434 0.02452 0.01737 -0.1248 0.6801 0.0221
-4.250 0.1701 0.02308 0.01567 -0.1251 0.6783 0.0225
-4.000 0.1973 0.02191 0.01423 -0.1254 0.6768 0.0230
-3.750 0.2249 0.02086 0.01292 -0.1255 0.6754 0.0237
-3.500 0.2524 0.02001 0.01189 -0.1257 0.6728 0.0247
-3.250 0.2803 0.01929 0.01091 -0.1258 0.6702 0.0264
-3.000 0.3076 0.01862 0.01022 -0.1259 0.6679 0.0275
-2.750 0.3353 0.01805 0.00957 -0.1260 0.6658 0.0286
-2.500 0.3631 0.01751 0.00894 -0.1260 0.6637 0.0299
-2.250 0.3912 0.01703 0.00833 -0.1259 0.6619 0.0316
-2.000 0.4191 0.01665 0.00795 -0.1261 0.6604 0.0344
-1.750 0.4469 0.01640 0.00765 -0.1261 0.6583 0.0377
-1.500 0.4742 0.01614 0.00744 -0.1263 0.6553 0.0410
-1.250 0.5019 0.01594 0.00723 -0.1264 0.6526 0.0473
-1.000 0.5298 0.01573 0.00702 -0.1264 0.6502 0.0566
-0.750 0.5579 0.01549 0.00682 -0.1266 0.6482 0.0761
-0.500 0.5860 0.01518 0.00666 -0.1267 0.6463 0.1263
-0.250 0.6138 0.01477 0.00656 -0.1270 0.6446 0.2373
0.000 0.6398 0.01449 0.00673 -0.1271 0.6417 0.3820
0.250 0.6645 0.01421 0.00693 -0.1268 0.6385 0.5340
0.500 0.6858 0.01371 0.00702 -0.1253 0.6357 0.7307
1.000 0.7510 0.01338 0.00680 -0.1268 0.6312 1.0000
1.250 0.7797 0.01342 0.00669 -0.1269 0.6293 1.0000
1.500 0.8061 0.01364 0.00689 -0.1269 0.6256 1.0000
1.750 0.8329 0.01381 0.00702 -0.1269 0.6220 1.0000
2.000 0.8605 0.01391 0.00705 -0.1269 0.6190 1.0000
2.250 0.8886 0.01395 0.00702 -0.1270 0.6165 1.0000
2.500 0.9172 0.01397 0.00696 -0.1271 0.6143 1.0000
2.750 0.9430 0.01421 0.00721 -0.1270 0.6105 1.0000
3.000 0.9691 0.01440 0.00740 -0.1269 0.6064 1.0000
3.250 0.9964 0.01448 0.00745 -0.1269 0.6033 1.0000
3.500 1.0246 0.01450 0.00742 -0.1269 0.6005 1.0000
3.750 1.0514 0.01462 0.00753 -0.1269 0.5971 1.0000
4.000 1.0762 0.01487 0.00782 -0.1267 0.5922 1.0000
4.250 1.1029 0.01495 0.00790 -0.1266 0.5883 1.0000
4.500 1.1309 0.01495 0.00787 -0.1266 0.5852 1.0000
4.750 1.1556 0.01518 0.00814 -0.1264 0.5803 1.0000
5.000 1.1806 0.01535 0.00836 -0.1261 0.5753 1.0000
5.250 1.2077 0.01538 0.00839 -0.1260 0.5713 1.0000
5.500 1.2328 0.01555 0.00859 -0.1258 0.5664 1.0000
5.750 1.2566 0.01576 0.00885 -0.1254 0.5605 1.0000
6.000 1.2835 0.01578 0.00885 -0.1252 0.5559 1.0000
6.250 1.3062 0.01606 0.00922 -0.1247 0.5495 1.0000
6.500 1.3306 0.01620 0.00938 -0.1243 0.5434 1.0000
6.750 1.3547 0.01636 0.00957 -0.1239 0.5372 1.0000
7.000 1.3769 0.01661 0.00988 -0.1232 0.5299 1.0000
7.250 1.4002 0.01678 0.01007 -0.1226 0.5229 1.0000
7.500 1.4209 0.01706 0.01040 -0.1217 0.5145 1.0000
7.750 1.4413 0.01735 0.01072 -0.1208 0.5061 1.0000
8.000 1.4615 0.01763 0.01101 -0.1198 0.4972 1.0000
8.250 1.4775 0.01807 0.01151 -0.1183 0.4872 1.0000
8.500 1.4938 0.01847 0.01193 -0.1168 0.4774 1.0000
8.750 1.5058 0.01895 0.01241 -0.1146 0.4666 1.0000
9.000 1.5129 0.01967 0.01317 -0.1119 0.4552 1.0000
9.250 1.5204 0.02049 0.01401 -0.1096 0.4434 1.0000
9.500 1.5275 0.02143 0.01494 -0.1074 0.4309 1.0000
9.750 1.5336 0.02251 0.01601 -0.1053 0.4175 1.0000
10.000 1.5387 0.02372 0.01721 -0.1032 0.4037 1.0000
10.250 1.5427 0.02507 0.01854 -0.1011 0.3892 1.0000
10.500 1.5458 0.02655 0.01999 -0.0991 0.3744 1.0000
10.750 1.5484 0.02813 0.02155 -0.0971 0.3595 1.0000
11.000 1.5499 0.02983 0.02323 -0.0952 0.3443 1.0000
11.250 1.5508 0.03165 0.02503 -0.0933 0.3287 1.0000
11.500 1.5510 0.03357 0.02692 -0.0915 0.3131 1.0000
11.750 1.5507 0.03558 0.02891 -0.0898 0.2977 1.0000
12.000 1.5495 0.03773 0.03102 -0.0881 0.2821 1.0000
12.250 1.5488 0.03993 0.03320 -0.0866 0.2671 1.0000
12.500 1.5480 0.04225 0.03550 -0.0852 0.2519 1.0000
12.750 1.5469 0.04469 0.03792 -0.0839 0.2370 1.0000
13.000 1.5455 0.04725 0.04046 -0.0828 0.2220 1.0000
13.250 1.5444 0.04985 0.04305 -0.0818 0.2077 1.0000
13.500 1.5430 0.05258 0.04576 -0.0808 0.1937 1.0000
13.750 1.5413 0.05542 0.04859 -0.0800 0.1801 1.0000
14.000 1.5398 0.05834 0.05149 -0.0793 0.1671 1.0000
14.250 1.5378 0.06136 0.05450 -0.0787 0.1544 1.0000
14.500 1.5354 0.06452 0.05765 -0.0783 0.1426 1.0000
14.750 1.5333 0.06772 0.06086 -0.0779 0.1307 1.0000
15.000 1.5319 0.07090 0.06406 -0.0776 0.1196 1.0000
15.250 1.5302 0.07419 0.06737 -0.0775 0.1095 1.0000
15.500 1.5279 0.07762 0.07081 -0.0774 0.1002 1.0000
15.750 1.5248 0.08124 0.07445 -0.0775 0.0915 1.0000
16.000 1.5224 0.08482 0.07807 -0.0777 0.0832 1.0000
16.250 1.5207 0.08834 0.08164 -0.0779 0.0761 1.0000
16.500 1.5168 0.09226 0.08560 -0.0783 0.0697 1.0000
16.750 1.5151 0.09592 0.08933 -0.0788 0.0637 1.0000
17.000 1.5119 0.09985 0.09331 -0.0794 0.0586 1.0000
17.250 1.5093 0.10373 0.09726 -0.0801 0.0537 1.0000
17.500 1.5056 0.10781 0.10141 -0.0810 0.0499 1.0000
17.750 1.5035 0.11173 0.10541 -0.0819 0.0459 1.0000
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