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 = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.3315 0.09086 0.08669 -0.0425 1.0000 0.0207 -7.750 -0.3338 0.08816 0.08408 -0.0432 1.0000 0.0207 -7.500 -0.3419 0.08540 0.08144 -0.0441 1.0000 0.0209 -7.250 -0.3502 0.08287 0.07901 -0.0441 1.0000 0.0209 -7.000 -0.3555 0.08010 0.07631 -0.0447 0.9997 0.0210 -6.750 -0.3322 0.07479 0.07085 -0.0539 0.9813 0.0214 -6.500 -0.3111 0.07082 0.06666 -0.0598 0.9623 0.0216 -6.000 -0.2784 0.05923 0.05502 -0.0664 0.9314 0.0242 -5.750 -0.2572 0.05490 0.05048 -0.0697 0.9118 0.0265 -5.500 -0.2313 0.05108 0.04626 -0.0727 0.8915 0.0300 -5.250 -0.2107 0.04734 0.04216 -0.0741 0.8678 0.0327 -5.000 -0.1863 0.04728 0.04130 -0.0740 0.8411 0.0404 -4.750 -0.1717 0.04054 0.03473 -0.0752 0.8181 0.0457 -3.750 -0.1018 0.02996 0.02268 -0.0734 0.7352 0.1452 -3.250 -0.0398 0.02645 0.01792 -0.0710 0.7053 0.1026 -3.000 -0.0032 0.02532 0.01620 -0.0693 0.6916 0.0565 -2.750 0.0281 0.02420 0.01466 -0.0680 0.6785 0.0364 -2.500 0.0569 0.02283 0.01297 -0.0670 0.6659 0.0269 -2.250 0.0859 0.02247 0.01217 -0.0659 0.6543 0.0222 -2.000 0.1129 0.02109 0.01063 -0.0654 0.6430 0.0212 -1.750 0.1389 0.01979 0.00923 -0.0648 0.6326 0.0205 -1.500 0.1628 0.01866 0.00805 -0.0637 0.6234 0.0193 -1.250 0.1854 0.01792 0.00715 -0.0626 0.6142 0.0182 -1.000 0.2091 0.01738 0.00636 -0.0617 0.6054 0.0175 -0.500 0.2597 0.01667 0.00486 -0.0604 0.5899 0.0204 -0.250 0.3662 0.01363 0.00428 -0.0760 0.5808 1.0000 0.000 0.3888 0.01374 0.00413 -0.0753 0.5740 1.0000 0.250 0.4117 0.01387 0.00397 -0.0747 0.5683 1.0000 0.500 0.4347 0.01401 0.00397 -0.0741 0.5619 1.0000 0.750 0.4582 0.01417 0.00398 -0.0736 0.5565 1.0000 1.000 0.4814 0.01434 0.00409 -0.0731 0.5509 1.0000 1.500 0.5291 0.01472 0.00444 -0.0723 0.5417 1.0000 1.750 0.5528 0.01493 0.00474 -0.0718 0.5371 1.0000 2.000 0.5717 0.01495 0.00474 -0.0701 0.5148 1.0000 2.250 0.5776 0.01498 0.00425 -0.0658 0.4182 1.0000 2.750 0.5933 0.01808 0.00573 -0.0596 0.0199 1.0000 3.000 0.6140 0.01856 0.00660 -0.0583 0.0172 1.0000 3.250 0.6338 0.01909 0.00739 -0.0569 0.0169 1.0000 3.500 0.6522 0.01980 0.00838 -0.0551 0.0169 1.0000 3.750 0.6687 0.02061 0.00948 -0.0531 0.0172 1.0000 4.000 0.6832 0.02155 0.01068 -0.0508 0.0171 1.0000 4.250 0.6950 0.02259 0.01202 -0.0483 0.0167 1.0000 4.500 0.7032 0.02376 0.01341 -0.0451 0.0162 1.0000 4.750 0.7086 0.02495 0.01475 -0.0417 0.0167 1.0000 5.000 0.7107 0.02624 0.01610 -0.0377 0.0177 1.0000 5.250 0.7157 0.02777 0.01745 -0.0341 0.0197 1.0000 5.500 0.7384 0.02823 0.01808 -0.0323 0.0245 1.0000 5.750 0.7893 0.03050 0.01999 -0.0342 0.0353 1.0000 7.750 1.0078 0.04916 0.04025 -0.0278 0.0325 1.0000 8.250 1.0228 0.05471 0.04645 -0.0229 0.0244 1.0000 8.500 1.0285 0.05758 0.04963 -0.0204 0.0219 1.0000 8.750 1.0466 0.06964 0.06144 -0.0229 0.0192 1.0000 9.000 1.0379 0.06936 0.06167 -0.0181 0.0189 1.0000 9.250 1.0301 0.06915 0.06191 -0.0136 0.0184 1.0000 9.500 1.0234 0.07020 0.06329 -0.0098 0.0176 1.0000 9.750 1.0162 0.07250 0.06582 -0.0068 0.0174 1.0000 10.000 1.0027 0.07457 0.06810 -0.0033 0.0167 1.0000 10.250 0.9891 0.07721 0.07093 -0.0008 0.0164 1.0000 10.500 0.9755 0.08021 0.07409 0.0007 0.0164 1.0000 10.750 0.9600 0.08345 0.07750 0.0016 0.0162 1.0000 11.000 0.9429 0.08715 0.08136 0.0017 0.0159 1.0000 11.250 0.9265 0.09116 0.08552 0.0010 0.0158 1.0000 11.500 0.9124 0.09547 0.08996 -0.0002 0.0164 1.0000 11.750 0.8949 0.10032 0.09494 -0.0021 0.0162 1.0000 12.000 0.8795 0.10546 0.10019 -0.0045 0.0164 1.0000