Eiffel 10 (Wright) - 1903 Wright Flyer airfoil (eiffel10-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: Eiffel 10 (Wright) - 1903 Wright Flyer airfoil (eiffel10-il) Reynolds number: 100,000 Max Cl/Cd: 29.88 at α=2.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-eiffel10-il-100000-n5.txt Download as CSV file: xf-eiffel10-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: Eiffel 10 (Wright) - 1903 Wright Flyer 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3907 0.11113 0.10566 -0.0095 1.0000 0.0218
-8.500 -0.3828 0.10873 0.10330 -0.0111 1.0000 0.0218
-8.000 -0.3667 0.10303 0.09769 -0.0131 1.0000 0.0221
-7.750 -0.3577 0.10017 0.09489 -0.0142 1.0000 0.0224
-7.500 -0.3471 0.09740 0.09216 -0.0159 1.0000 0.0228
-7.250 -0.3357 0.09468 0.08949 -0.0178 1.0000 0.0231
-7.000 -0.3235 0.09198 0.08684 -0.0199 1.0000 0.0236
-6.750 -0.3106 0.08931 0.08422 -0.0222 1.0000 0.0240
-6.500 -0.2968 0.08670 0.08167 -0.0247 1.0000 0.0245
-6.250 -0.2812 0.08431 0.07932 -0.0280 1.0000 0.0250
-5.750 -0.2586 0.07919 0.07434 -0.0307 1.0000 0.0259
-5.500 -0.2489 0.07694 0.07216 -0.0315 1.0000 0.0267
-5.250 -0.2384 0.07485 0.07012 -0.0326 1.0000 0.0276
-5.000 -0.2104 0.07203 0.06730 -0.0382 0.9970 0.0287
-4.750 -0.1688 0.06831 0.06355 -0.0466 0.9915 0.0293
-4.500 -0.1337 0.06444 0.05969 -0.0520 0.9849 0.0307
-4.250 -0.0595 0.06000 0.05510 -0.0672 0.9722 0.0333
-4.000 -0.0148 0.05550 0.05058 -0.0737 0.9475 0.0349
-3.750 0.0345 0.05236 0.04734 -0.0816 0.9307 0.0380
-3.500 0.0760 0.04902 0.04392 -0.0871 0.9062 0.0395
-3.250 0.1101 0.04638 0.04122 -0.0905 0.8862 0.0425
-3.000 0.1531 0.04394 0.03846 -0.0951 0.8363 0.0452
-2.750 0.1768 0.04164 0.03603 -0.0954 0.7961 0.0482
-2.500 0.2162 0.03968 0.03368 -0.0988 0.7358 0.0528
-2.000 0.2646 0.03899 0.03012 -0.1003 0.0476 0.0632
-1.750 0.3045 0.03733 0.02827 -0.1036 0.0437 0.0720
-1.500 0.3423 0.03601 0.02680 -0.1062 0.0415 0.0832
-1.000 0.4025 0.03270 0.02352 -0.1088 0.0373 0.1037
0.000 0.5488 0.02737 0.01730 -0.1137 0.0330 0.0580
0.250 0.5829 0.02625 0.01586 -0.1139 0.0328 0.0435
0.500 0.6126 0.02573 0.01519 -0.1139 0.0327 0.0421
0.750 0.6416 0.02540 0.01465 -0.1136 0.0327 0.0410
1.000 0.6700 0.02528 0.01419 -0.1130 0.0328 0.0383
1.250 0.6964 0.02528 0.01405 -0.1123 0.0330 0.0372
1.500 0.7232 0.02545 0.01405 -0.1115 0.0333 0.0362
1.750 0.7502 0.02576 0.01425 -0.1107 0.0337 0.0357
2.000 0.7775 0.02626 0.01467 -0.1100 0.0342 0.0356
2.250 0.8043 0.02694 0.01527 -0.1091 0.0349 0.0372
2.500 0.8313 0.02782 0.01608 -0.1082 0.0357 0.0395
2.750 0.8587 0.02893 0.01713 -0.1074 0.0365 0.0410
3.000 0.8862 0.03035 0.01846 -0.1068 0.0373 0.0429
3.250 0.9135 0.03172 0.01984 -0.1061 0.0375 0.0464
3.500 0.9405 0.03281 0.02114 -0.1052 0.0369 0.0582
3.750 0.9638 0.03290 0.02277 -0.1040 0.0366 1.0000
4.000 0.9898 0.03521 0.02499 -0.1033 0.0376 1.0000
4.250 1.0178 0.03663 0.02686 -0.1016 0.0425 1.0000
4.500 1.0427 0.04013 0.03023 -0.1012 0.0452 1.0000
5.000 1.1318 0.04425 0.03663 -0.0948 0.1639 1.0000
5.500 1.1694 0.05013 0.04271 -0.0929 0.1430 1.0000
6.000 1.1956 0.05635 0.04914 -0.0910 0.1121 1.0000
7.500 1.2392 0.07791 0.07189 -0.0839 0.0716 1.0000
7.750 1.2431 0.08164 0.07573 -0.0828 0.0691 1.0000
8.000 1.2512 0.08590 0.07995 -0.0821 0.0673 1.0000
8.250 1.2719 0.09525 0.08890 -0.0835 0.0655 1.0000
8.500 1.2488 0.09728 0.09147 -0.0809 0.0651 1.0000
8.750 1.2186 0.10025 0.09484 -0.0792 0.0644 1.0000
9.000 1.1936 0.10435 0.09908 -0.0783 0.0642 1.0000
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