XFOIL Version 6.96 Calculated polar for: WORTMANN FX 049-915 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 ------ -------- --------- --------- -------- -------- -------- -12.500 -0.3610 0.08069 0.07679 -0.0715 0.4685 0.0074 -12.250 -0.6790 0.03802 0.03376 -0.0893 0.4735 0.0067 -12.000 -0.7044 0.03169 0.02705 -0.0860 0.4732 0.0066 -11.750 -0.7069 0.02767 0.02275 -0.0836 0.4724 0.0067 -11.500 -0.6962 0.02548 0.02038 -0.0820 0.4714 0.0068 -11.250 -0.6829 0.02386 0.01859 -0.0803 0.4703 0.0069 -11.000 -0.6650 0.02277 0.01737 -0.0791 0.4693 0.0071 -10.750 -0.6474 0.02167 0.01612 -0.0778 0.4681 0.0072 -10.500 -0.6289 0.02064 0.01494 -0.0765 0.4670 0.0074 -10.250 -0.6085 0.01982 0.01400 -0.0754 0.4659 0.0075 -10.000 -0.5879 0.01904 0.01307 -0.0743 0.4647 0.0076 -9.750 -0.5665 0.01833 0.01224 -0.0732 0.4635 0.0077 -9.500 -0.5448 0.01764 0.01143 -0.0722 0.4624 0.0078 -9.250 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0.4180 1.0000 5.750 1.1761 0.00866 0.00375 -0.0786 0.4171 1.0000 6.000 1.2007 0.00877 0.00388 -0.0781 0.4161 1.0000 6.250 1.2251 0.00891 0.00401 -0.0776 0.4144 1.0000 6.500 1.2503 0.00894 0.00410 -0.0773 0.4131 1.0000 6.750 1.2757 0.00895 0.00417 -0.0770 0.4107 1.0000 7.000 1.3010 0.00900 0.00425 -0.0766 0.4083 1.0000 7.250 1.3260 0.00905 0.00434 -0.0763 0.4057 1.0000 7.500 1.3502 0.00911 0.00438 -0.0758 0.4012 1.0000 7.750 1.3753 0.00916 0.00448 -0.0755 0.3964 1.0000 8.000 1.4001 0.00923 0.00458 -0.0751 0.3910 1.0000 8.250 1.4237 0.00935 0.00470 -0.0745 0.3849 1.0000 8.500 1.4471 0.00948 0.00483 -0.0740 0.3730 1.0000 8.750 1.4646 0.00993 0.00515 -0.0724 0.3441 1.0000 9.000 1.4691 0.01114 0.00608 -0.0689 0.2937 1.0000 9.250 1.4711 0.01245 0.00717 -0.0650 0.2524 1.0000 9.500 1.4683 0.01389 0.00839 -0.0604 0.2107 1.0000 9.750 1.4480 0.01595 0.01018 -0.0533 0.1603 1.0000 10.000 1.4195 0.01776 0.01187 -0.0451 0.1312 1.0000 10.250 1.3965 0.02042 0.01452 -0.0406 0.1137 1.0000 10.500 1.3778 0.02499 0.01917 -0.0406 0.1038 1.0000 10.750 1.3704 0.02889 0.02315 -0.0410 0.0992 1.0000 11.000 1.3443 0.03417 0.02846 -0.0403 0.0895 1.0000 11.250 1.3340 0.03758 0.03194 -0.0396 0.0862 1.0000 11.500 1.3179 0.04145 0.03583 -0.0386 0.0799 1.0000 11.750 1.3063 0.04489 0.03929 -0.0378 0.0742 1.0000 12.000 1.2975 0.04807 0.04248 -0.0371 0.0671 1.0000 12.250 1.2840 0.05183 0.04621 -0.0365 0.0549 1.0000 12.500 1.2632 0.05654 0.05079 -0.0360 0.0320 1.0000 12.750 1.2590 0.05965 0.05391 -0.0357 0.0250 1.0000 13.000 1.2591 0.06232 0.05657 -0.0355 0.0188 1.0000 13.250 1.2617 0.06477 0.05906 -0.0353 0.0169 1.0000 13.500 1.2636 0.06733 0.06164 -0.0352 0.0149 1.0000 13.750 1.2687 0.06952 0.06387 -0.0352 0.0139 1.0000 14.000 1.2728 0.07184 0.06620 -0.0351 0.0126 1.0000 14.250 1.2756 0.07436 0.06877 -0.0351 0.0116 1.0000 14.500 1.2797 0.07675 0.07120 -0.0351 0.0109 1.0000 14.750 1.2846 0.07901 0.07348 -0.0352 0.0100 1.0000 15.000 1.2900 0.08122 0.07573 -0.0352 0.0097 1.0000 15.250 1.2933 0.08374 0.07828 -0.0353 0.0091 1.0000 15.500 1.2962 0.08635 0.08093 -0.0355 0.0084 1.0000 15.750 1.3015 0.08864 0.08325 -0.0356 0.0080 1.0000 16.000 1.3063 0.09096 0.08561 -0.0358 0.0077 1.0000 16.250 1.3095 0.09353 0.08822 -0.0360 0.0073 1.0000 16.500 1.3121 0.09622 0.09095 -0.0363 0.0069 1.0000