XFOIL Version 6.94 Calculated polar for: rt 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2974 0.04477 0.03335 0.0508 0.9991 0.6042 -2.750 -0.2748 0.04550 0.03401 0.0501 0.9991 0.6116 -2.500 -0.1731 0.04658 0.03512 0.0343 0.9687 0.6292 -2.250 -0.0764 0.04644 0.03501 0.0206 0.9340 0.6483 -2.000 0.0171 0.04525 0.03395 0.0088 0.8990 0.6699 -1.750 0.1117 0.04274 0.03166 -0.0018 0.8588 0.6931 -1.500 0.2086 0.03914 0.02852 -0.0115 0.8188 0.7257 -1.250 0.3654 0.03140 0.02160 -0.0282 0.7268 0.8009 -1.000 0.5068 0.02974 0.01894 -0.0472 0.5238 0.8885 -0.750 0.5754 0.03217 0.02046 -0.0578 0.4721 1.0009 -0.500 0.6054 0.03280 0.02052 -0.0591 0.4522 1.0009 -0.250 0.6380 0.03347 0.02067 -0.0606 0.4371 1.0009 0.000 0.6706 0.03414 0.02108 -0.0620 0.4253 1.0009 0.250 0.7051 0.03483 0.02144 -0.0638 0.4158 1.0009 0.500 0.7365 0.03557 0.02210 -0.0649 0.4081 1.0009 0.750 0.7692 0.03634 0.02263 -0.0662 0.4004 1.0009 1.000 0.7999 0.03721 0.02337 -0.0671 0.3950 1.0009 1.250 0.8268 0.03820 0.02440 -0.0674 0.3905 1.0009 1.500 0.8536 0.03924 0.02549 -0.0677 0.3869 1.0009 1.750 0.8802 0.04038 0.02664 -0.0679 0.3840 1.0009 2.000 0.9058 0.04161 0.02789 -0.0680 0.3819 1.0009 2.250 0.9306 0.04295 0.02926 -0.0679 0.3802 1.0009 2.500 0.9541 0.04443 0.03083 -0.0677 0.3789 1.0009 2.750 0.9757 0.04608 0.03258 -0.0673 0.3781 1.0009 3.000 0.9951 0.04789 0.03451 -0.0665 0.3776 1.0009 3.250 1.0112 0.04993 0.03673 -0.0653 0.3776 1.0009 3.500 1.0234 0.05226 0.03931 -0.0636 0.3778 1.0009 3.750 1.0280 0.05511 0.04245 -0.0611 0.3788 1.0009 4.000 1.0184 0.05906 0.04677 -0.0570 0.3804 1.0009 4.250 0.9710 0.06605 0.05423 -0.0494 0.3842 1.0009 4.500 0.8657 0.07840 0.06687 -0.0387 0.3920 1.0009 4.750 0.8101 0.08805 0.07659 -0.0351 0.3987 1.0009