XFOIL Version 6.96 Calculated polar for: GOE 559 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 ------ -------- --------- --------- -------- -------- -------- -1.750 0.1455 0.01339 0.00678 -0.0651 0.5025 0.0060 -1.500 0.1727 0.01351 0.00666 -0.0645 0.4971 0.0066 -1.250 0.1974 0.01132 0.00452 -0.0640 0.4927 0.0081 -1.000 0.2227 0.01071 0.00381 -0.0635 0.4879 0.0110 -0.750 0.2466 0.01005 0.00308 -0.0626 0.4835 0.0146 -0.500 0.2702 0.00942 0.00246 -0.0619 0.4799 0.0204 -0.250 0.2957 0.00909 0.00202 -0.0612 0.4759 0.0101 0.000 0.3211 0.00882 0.00162 -0.0606 0.4721 0.0051 0.250 0.3470 0.00863 0.00131 -0.0603 0.4686 0.0046 0.750 0.4227 0.00621 0.00144 -0.0658 0.4616 0.9823 1.000 0.4576 0.00631 0.00148 -0.0677 0.4579 0.9859 1.250 0.4918 0.00640 0.00154 -0.0695 0.4540 0.9891 1.750 0.5451 0.00899 0.00256 -0.0712 0.0042 0.9964 2.000 0.5795 0.00937 0.00305 -0.0730 0.0053 0.9986 2.250 0.6142 0.00947 0.00323 -0.0744 0.0224 1.0000 2.500 0.6314 0.01004 0.00391 -0.0724 0.0181 1.0000 2.750 0.6474 0.01068 0.00467 -0.0701 0.0136 1.0000 3.000 0.6470 0.01233 0.00641 -0.0649 0.0109 1.0000 3.250 0.6739 0.01220 0.00634 -0.0645 0.0089 1.0000 3.500 0.6859 0.01301 0.00719 -0.0615 0.0071 1.0000 3.750 0.6885 0.01459 0.00871 -0.0570 0.0063 1.0000 4.000 0.7024 0.01585 0.01000 -0.0543 0.0059 1.0000 4.250 0.7214 0.01712 0.01124 -0.0527 0.0033 1.0000 4.500 0.7449 0.01711 0.01137 -0.0511 0.0029 1.0000 4.750 0.7697 0.01824 0.01252 -0.0500 0.0023 1.0000 5.000 0.7959 0.01982 0.01411 -0.0495 0.0019 1.0000 5.250 0.8269 0.02351 0.01777 -0.0503 0.0016 1.0000 5.500 0.8527 0.02512 0.01949 -0.0492 0.0016 1.0000 5.750 0.8781 0.02645 0.02095 -0.0478 0.0015 1.0000 6.000 0.9054 0.02732 0.02194 -0.0462 0.0013 1.0000 6.250 0.9289 0.02936 0.02411 -0.0446 0.0012 1.0000 6.500 0.9496 0.03140 0.02629 -0.0428 0.0011 1.0000 6.750 0.9684 0.03367 0.02873 -0.0406 0.0010 1.0000 7.000 0.9857 0.03606 0.03130 -0.0383 0.0009 1.0000 7.250 1.0008 0.03852 0.03394 -0.0358 0.0008 1.0000 7.500 1.0142 0.04104 0.03665 -0.0332 0.0008 1.0000 7.750 1.0259 0.04365 0.03945 -0.0306 0.0008 1.0000 8.000 1.0350 0.04646 0.04245 -0.0278 0.0007 1.0000 8.250 1.0412 0.04927 0.04544 -0.0252 0.0007 1.0000 8.500 1.0459 0.05220 0.04855 -0.0223 0.0007 1.0000 8.750 1.0483 0.05520 0.05173 -0.0192 0.0007 1.0000 9.000 1.0477 0.05826 0.05495 -0.0162 0.0006 1.0000 9.250 1.0424 0.06159 0.05843 -0.0135 0.0006 1.0000 9.500 1.0347 0.06500 0.06198 -0.0105 0.0006 1.0000 9.750 1.0221 0.06932 0.06641 -0.0081 0.0006 1.0000 10.000 1.0083 0.07187 0.06907 -0.0042 0.0006 1.0000 10.250 0.9918 0.07443 0.07174 -0.0008 0.0006 1.0000 10.500 0.9738 0.07752 0.07493 0.0015 0.0006 1.0000 10.750 0.9583 0.08016 0.07767 0.0029 0.0006 1.0000 11.000 0.9451 0.08191 0.07957 0.0040 0.0006 1.0000 11.250 0.9289 0.08530 0.08308 0.0037 0.0006 1.0000 11.500 0.9103 0.09016 0.08800 0.0024 0.0006 1.0000 11.750 0.8941 0.09417 0.09214 0.0008 0.0006 1.0000