XFOIL Version 6.96 Calculated polar for: EPPLER E850 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.4385 0.09119 0.08970 -0.0267 1.0000 0.0056 -9.250 -0.4441 0.08738 0.08591 -0.0267 1.0000 0.0056 -9.000 -0.4507 0.08358 0.08212 -0.0266 1.0000 0.0057 -8.750 -0.4499 0.07815 0.07671 -0.0295 0.9994 0.0059 -8.500 -0.4495 0.07233 0.07089 -0.0331 0.9984 0.0059 -8.250 -0.4500 0.06602 0.06458 -0.0374 0.9972 0.0059 -8.000 -0.4531 0.05849 0.05706 -0.0436 0.9956 0.0058 -7.750 -0.4728 0.04663 0.04511 -0.0588 0.9913 0.0051 -7.500 -0.4714 0.04019 0.03850 -0.0649 0.9870 0.0052 -7.250 -0.4619 0.03469 0.03283 -0.0693 0.9845 0.0054 -7.000 -0.4465 0.02972 0.02766 -0.0729 0.9828 0.0057 -6.750 -0.4272 0.02518 0.02291 -0.0759 0.9817 0.0062 -6.500 -0.4139 0.02176 0.01928 -0.0762 0.9779 0.0068 -6.250 -0.3951 0.01844 0.01571 -0.0767 0.9748 0.0077 -6.000 -0.3701 0.01583 0.01287 -0.0775 0.9731 0.0091 -5.750 -0.3367 0.01625 0.01313 -0.0777 0.9722 0.0108 -5.500 -0.3105 0.01354 0.01011 -0.0788 0.9710 0.0109 -5.250 -0.3134 0.01934 0.01512 -0.0798 0.9704 0.0124 -5.000 -0.2838 0.01741 0.01305 -0.0810 0.9697 0.0144 -4.750 -0.2525 0.01602 0.01149 -0.0820 0.9692 0.0173 -4.500 -0.2191 0.01542 0.01077 -0.0829 0.9688 0.0212 -4.250 -1.2480 0.06364 0.06202 0.2358 0.9815 0.0050 -4.000 -0.1756 0.01028 0.00521 -0.0786 0.9622 0.0070 -3.750 -0.1452 0.00933 0.00409 -0.0790 0.9609 0.0048 -3.500 -0.1134 0.00891 0.00360 -0.0798 0.9600 0.0041 -3.250 -0.0812 0.00833 0.00283 -0.0808 0.9592 0.0038 -3.000 -0.0482 0.00809 0.00242 -0.0819 0.9586 0.0037 -2.750 -0.0149 0.00791 0.00220 -0.0831 0.9580 0.0040 -2.500 0.0178 0.00765 0.00202 -0.0844 0.9574 0.0335 -2.250 0.0492 0.00691 0.00179 -0.0858 0.9567 0.1845 -2.000 0.0787 0.00572 0.00165 -0.0873 0.9559 0.4918 -1.750 0.1011 0.00556 0.00171 -0.0862 0.9519 0.5611 -1.500 0.1307 0.00547 0.00165 -0.0866 0.9496 0.5755 -1.250 0.1626 0.00536 0.00156 -0.0876 0.9476 0.5883 -1.000 0.1962 0.00523 0.00145 -0.0888 0.9456 0.5989 -0.750 0.2300 0.00509 0.00134 -0.0902 0.9434 0.6075 -0.500 0.2563 0.00502 0.00130 -0.0898 0.9382 0.6159 -0.250 0.2870 0.00491 0.00121 -0.0904 0.9335 0.6251 0.000 0.3178 0.00481 0.00114 -0.0910 0.9285 0.6343 0.250 0.3449 0.00473 0.00110 -0.0908 0.9208 0.6440 0.500 0.3732 0.00467 0.00106 -0.0908 0.9125 0.6544 1.000 0.4269 0.00456 0.00100 -0.0901 0.8830 0.6774 1.250 0.4529 0.00454 0.00098 -0.0895 0.8599 0.6898 1.500 0.4803 0.00454 0.00111 -0.0894 0.8520 0.7034 2.000 0.5178 0.00518 0.00123 -0.0851 0.6782 0.7331 2.250 0.5299 0.00604 0.00152 -0.0819 0.5217 0.7505 2.500 0.5434 0.00709 0.00187 -0.0794 0.3315 0.7701 2.750 0.5603 0.00801 0.00223 -0.0775 0.1718 0.7920 3.000 0.5806 0.00865 0.00256 -0.0763 0.0758 0.8173 3.250 0.5995 0.00956 0.00334 -0.0742 0.0063 0.8476 3.500 0.6197 0.01007 0.00415 -0.0723 0.0055 0.8886 3.750 0.6438 0.01009 0.00433 -0.0714 0.0046 0.9752 4.000 0.6694 0.01070 0.00503 -0.0709 0.0040 1.0000 4.250 0.6936 0.01122 0.00560 -0.0702 0.0034 1.0000 4.500 0.7181 0.01168 0.00607 -0.0696 0.0026 1.0000 4.750 0.7362 0.01392 0.00854 -0.0673 0.0020 1.0000 5.750 0.8012 0.03377 0.03031 -0.0544 0.0057 1.0000 6.000 0.8311 0.03331 0.02990 -0.0540 0.0047 1.0000 6.250 0.8486 0.03612 0.03292 -0.0517 0.0042 1.0000 6.500 0.8615 0.03981 0.03686 -0.0490 0.0039 1.0000 6.750 0.8717 0.04381 0.04111 -0.0462 0.0037 1.0000 7.000 0.8797 0.04794 0.04546 -0.0436 0.0035 1.0000 7.250 0.8850 0.05227 0.05000 -0.0410 0.0034 1.0000 7.500 0.8876 0.05670 0.05461 -0.0385 0.0033 1.0000 7.750 0.8874 0.06113 0.05920 -0.0363 0.0032 1.0000 8.000 0.8833 0.06570 0.06392 -0.0343 0.0032 1.0000 8.250 0.8768 0.06994 0.06828 -0.0326 0.0031 1.0000 8.500 0.8629 0.07390 0.07232 -0.0303 0.0031 1.0000 8.750 0.8448 0.07776 0.07626 -0.0284 0.0032 1.0000 9.000 0.8268 0.08238 0.08093 -0.0289 0.0032 1.0000