XFOIL Version 6.96 Calculated polar for: GOE 5K 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.750 -0.5247 0.08940 0.08782 -0.0098 1.0000 0.0017 -7.500 -0.5285 0.08682 0.08527 -0.0096 1.0000 0.0017 -7.250 -0.5279 0.08351 0.08198 -0.0112 1.0000 0.0017 -7.000 -0.5237 0.07993 0.07840 -0.0137 1.0000 0.0017 -6.750 -0.5176 0.07615 0.07462 -0.0163 1.0000 0.0017 -6.500 -0.5098 0.07235 0.07081 -0.0187 1.0000 0.0017 -6.250 -0.5006 0.06847 0.06691 -0.0210 1.0000 0.0017 -6.000 -0.4773 0.06334 0.06171 -0.0265 0.9990 0.0017 -5.750 -0.4498 0.05808 0.05637 -0.0323 0.9975 0.0018 -5.500 -0.4205 0.05293 0.05112 -0.0374 0.9960 0.0018 -5.250 -0.3956 0.04781 0.04589 -0.0415 0.9942 0.0018 -2.750 -0.1054 0.01329 0.00897 -0.0509 0.9738 0.0016 -2.500 -0.0764 0.01118 0.00658 -0.0505 0.9719 0.0014 -2.250 -0.0472 0.00924 0.00436 -0.0501 0.9705 0.0013 -2.000 -0.0185 0.00775 0.00263 -0.0499 0.9692 0.0015 -1.750 0.0110 0.00726 0.00203 -0.0503 0.9676 0.0028 -1.500 0.0345 0.00671 0.00137 -0.0492 0.9622 0.0033 -1.250 0.0640 0.00640 0.00101 -0.0496 0.9594 0.0064 -1.000 0.0944 0.00604 0.00076 -0.0504 0.9572 0.0588 -0.750 0.1188 0.00497 0.00066 -0.0504 0.9529 0.4085 -0.500 0.1388 0.00395 0.00063 -0.0491 0.9461 0.7378 0.000 0.2367 0.00334 0.00059 -0.0589 0.9341 0.9785 0.250 0.2943 0.00336 0.00055 -0.0659 0.9125 0.9981 0.500 0.3288 0.00346 0.00050 -0.0674 0.8702 1.0000 0.750 0.3490 0.00367 0.00049 -0.0655 0.8053 1.0000 1.000 0.3663 0.00403 0.00056 -0.0631 0.7217 1.0000 1.250 0.3683 0.00569 0.00082 -0.0580 0.3474 1.0000 1.500 0.3803 0.00714 0.00115 -0.0552 0.0204 1.0000 1.750 0.4034 0.00747 0.00146 -0.0542 0.0043 1.0000 2.000 0.4262 0.00785 0.00192 -0.0530 0.0021 1.0000 2.250 0.4484 0.00839 0.00256 -0.0516 0.0018 1.0000 2.500 0.4692 0.00916 0.00345 -0.0499 0.0018 1.0000 2.750 0.4895 0.01021 0.00464 -0.0480 0.0018 1.0000 3.000 0.5106 0.01175 0.00634 -0.0461 0.0020 1.0000 3.250 0.5355 0.01375 0.00856 -0.0442 0.0028 1.0000 3.500 0.5593 0.01505 0.01005 -0.0430 0.0014 1.0000 4.000 0.6042 0.01995 0.01552 -0.0390 0.0010 1.0000 4.500 0.6458 0.02644 0.02264 -0.0334 0.0014 1.0000 4.750 0.6645 0.02964 0.02612 -0.0308 0.0015 1.0000 5.000 0.6823 0.03293 0.02967 -0.0284 0.0014 1.0000 5.250 0.6991 0.03635 0.03332 -0.0262 0.0013 1.0000 5.500 0.7150 0.03995 0.03714 -0.0241 0.0013 1.0000 5.750 0.7290 0.04364 0.04103 -0.0224 0.0012 1.0000 6.000 0.7417 0.04741 0.04498 -0.0209 0.0012 1.0000 6.250 0.7512 0.05133 0.04906 -0.0197 0.0011 1.0000 6.500 0.7586 0.05605 0.05394 -0.0185 0.0011 1.0000 6.750 0.7674 0.06040 0.05842 -0.0178 0.0011 1.0000 7.000 0.7745 0.06466 0.06280 -0.0174 0.0011 1.0000 7.250 0.7795 0.06908 0.06731 -0.0173 0.0011 1.0000 7.500 0.7820 0.07369 0.07201 -0.0178 0.0011 1.0000 7.750 0.7828 0.07801 0.07638 -0.0186 0.0011 1.0000 8.000 0.7800 0.08251 0.08093 -0.0199 0.0011 1.0000 8.250 0.7715 0.08649 0.08493 -0.0209 0.0011 1.0000 8.500 0.7639 0.09128 0.08972 -0.0253 0.0011 1.0000