Eiffel 10 (Wright) - 1903 Wright Flyer airfoil (eiffel10-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: Eiffel 10 (Wright) - 1903 Wright Flyer airfoil (eiffel10-il) Reynolds number: 200,000 Max Cl/Cd: 28.74 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-eiffel10-il-200000.txt Download as CSV file: xf-eiffel10-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.3732 0.13029 0.12646 -0.0026 1.0000 0.0183
-11.250 -0.3683 0.12766 0.12384 -0.0037 1.0000 0.0183
-8.750 -0.4196 0.11180 0.10789 -0.0010 1.0000 0.0184
-8.500 -0.4108 0.10839 0.10451 -0.0015 1.0000 0.0185
-8.250 -0.4022 0.10543 0.10158 -0.0025 1.0000 0.0187
-8.000 -0.3937 0.10264 0.09883 -0.0037 1.0000 0.0188
-7.750 -0.3851 0.09996 0.09618 -0.0049 1.0000 0.0190
-7.500 -0.3752 0.09725 0.09351 -0.0066 1.0000 0.0192
-7.250 -0.3629 0.09442 0.09072 -0.0088 1.0000 0.0194
-7.000 -0.3496 0.09157 0.08790 -0.0113 1.0000 0.0196
-6.750 -0.3351 0.08869 0.08506 -0.0140 1.0000 0.0199
-6.500 -0.3194 0.08581 0.08221 -0.0169 1.0000 0.0202
-6.250 -0.3023 0.08292 0.07936 -0.0201 1.0000 0.0205
-6.000 -0.2830 0.08010 0.07657 -0.0239 1.0000 0.0209
-5.750 -0.2670 0.06807 0.06504 -0.0173 0.9998 0.0212
-5.500 -0.2338 0.06184 0.05882 -0.0224 0.9948 0.0218
-5.250 -0.1871 0.05580 0.05275 -0.0322 0.9867 0.0226
-5.000 -0.1466 0.05088 0.04781 -0.0408 0.9795 0.0238
-4.750 -0.1008 0.04555 0.04244 -0.0502 0.9623 0.0247
-4.500 -0.1133 0.06187 0.05842 -0.0555 0.9865 0.0243
-4.250 -0.0810 0.05742 0.05401 -0.0590 0.9679 0.0250
-4.000 -0.0396 0.05398 0.05054 -0.0652 0.9474 0.0260
-3.750 0.0007 0.05094 0.04738 -0.0706 0.9134 0.0272
-3.250 0.0613 0.04594 0.04200 -0.0755 0.8234 0.0283
-3.000 0.0834 0.04411 0.03969 -0.0754 0.7206 0.0293
-2.750 0.1050 0.04499 0.03801 -0.0768 0.0574 0.0304
-2.500 0.1570 0.04280 0.03557 -0.0830 0.0500 0.0318
-2.250 0.1789 0.04042 0.03326 -0.0836 0.0475 0.0328
-2.000 0.2145 0.03848 0.03126 -0.0863 0.0462 0.0349
-1.750 0.2557 0.03642 0.02907 -0.0897 0.0456 0.0367
-1.500 0.2858 0.03474 0.02741 -0.0911 0.0454 0.0388
-1.250 0.3271 0.03307 0.02556 -0.0938 0.0455 0.0418
-1.000 0.3563 0.03169 0.02421 -0.0949 0.0460 0.0447
-0.750 0.3926 0.03035 0.02272 -0.0966 0.0468 0.0486
-0.500 0.4291 0.02965 0.02175 -0.0978 0.0478 0.0548
-0.250 0.4558 0.02873 0.02084 -0.0984 0.0493 0.0592
0.000 0.4864 0.02842 0.02029 -0.0990 0.0507 0.0668
0.250 0.5190 0.02733 0.01918 -0.0994 0.0537 0.0774
0.500 0.5488 0.02717 0.01891 -0.0995 0.0566 0.0911
0.750 0.5793 0.02734 0.01885 -0.0996 0.0575 0.1144
1.000 0.6092 0.02700 0.01856 -0.0994 0.0638 0.1449
1.250 0.6384 0.02711 0.01884 -0.0989 0.0791 0.1818
3.500 0.9241 0.03215 0.02365 -0.0901 0.1145 0.0455
3.750 0.9476 0.03398 0.02552 -0.0899 0.0989 0.0453
4.000 0.9709 0.03572 0.02735 -0.0895 0.0859 0.0466
4.250 0.9935 0.03778 0.02960 -0.0889 0.0763 0.0546
4.500 1.0132 0.03834 0.03153 -0.0878 0.0720 1.0000
4.750 1.0345 0.04051 0.03392 -0.0867 0.0652 1.0000
5.000 1.0565 0.04269 0.03608 -0.0861 0.0616 1.0000
5.500 1.0957 0.04764 0.04146 -0.0840 0.0541 1.0000
5.750 1.1146 0.05048 0.04429 -0.0834 0.0522 1.0000
6.000 1.1331 0.05528 0.04887 -0.0836 0.0511 1.0000
6.250 1.1450 0.05648 0.05088 -0.0808 0.0472 1.0000
6.500 1.1596 0.05937 0.05389 -0.0798 0.0454 1.0000
6.750 1.1739 0.06288 0.05735 -0.0794 0.0444 1.0000
7.000 1.1916 0.06945 0.06360 -0.0802 0.0436 1.0000
7.250 1.1843 0.07031 0.06554 -0.0761 0.0407 1.0000
7.500 1.1895 0.07386 0.06924 -0.0751 0.0395 1.0000
7.750 1.1935 0.07749 0.07295 -0.0741 0.0388 1.0000
8.000 1.1975 0.08128 0.07678 -0.0733 0.0383 1.0000
8.250 1.2067 0.08567 0.08107 -0.0729 0.0379 1.0000
8.500 1.2127 0.09145 0.08680 -0.0728 0.0375 1.0000
8.750 1.1904 0.09523 0.09093 -0.0711 0.0374 1.0000
9.000 1.1678 0.09913 0.09501 -0.0697 0.0374 1.0000
9.250 1.1466 0.10366 0.09967 -0.0701 0.0374 1.0000
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