XFOIL Version 6.96 Calculated polar for: GOE 114 (MVA MK.1) 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 ------ -------- --------- --------- -------- -------- -------- -9.250 -0.3121 0.09332 0.08916 -0.0465 1.0000 0.0759 -9.000 -0.2943 0.08978 0.08562 -0.0427 1.0000 0.0829 -8.500 -0.3825 0.09484 0.09037 -0.0437 1.0000 0.0734 -5.500 -0.2965 0.03558 0.02923 -0.0873 0.9486 0.0267 -5.250 -0.2610 0.03296 0.02585 -0.0884 0.9442 0.0198 -5.000 -0.2331 0.02983 0.02229 -0.0892 0.9388 0.0181 -4.750 -0.2026 0.02744 0.01947 -0.0898 0.9334 0.0167 -4.500 -0.1677 0.02534 0.01697 -0.0909 0.9300 0.0157 -4.250 -0.1401 0.02376 0.01517 -0.0907 0.9236 0.0162 -4.000 -0.1088 0.02231 0.01357 -0.0912 0.9188 0.0178 -3.750 -0.0777 0.02105 0.01217 -0.0916 0.9142 0.0183 -3.500 -0.0513 0.02012 0.01109 -0.0913 0.9070 0.0186 -3.250 -0.0176 0.01915 0.00980 -0.0923 0.9027 0.0196 -3.000 0.0088 0.01830 0.00861 -0.0921 0.8951 0.0232 -2.750 0.0417 0.01752 0.00745 -0.0929 0.8901 0.0287 -2.500 0.0677 0.01594 0.00685 -0.0933 0.8835 0.2623 -2.250 0.0975 0.01641 0.00727 -0.0934 0.8772 0.3626 -2.000 0.1250 0.01656 0.00730 -0.0933 0.8705 0.3958 -1.750 0.1547 0.01641 0.00691 -0.0936 0.8641 0.4046 -1.500 0.1842 0.01626 0.00663 -0.0940 0.8579 0.4123 -1.250 0.2127 0.01611 0.00638 -0.0942 0.8510 0.4195 -1.000 0.2421 0.01596 0.00612 -0.0945 0.8447 0.4253 -0.750 0.2704 0.01586 0.00593 -0.0947 0.8374 0.4306 -0.500 0.2994 0.01570 0.00578 -0.0949 0.8308 0.4356 -0.250 0.3274 0.01560 0.00567 -0.0950 0.8234 0.4420 0.000 0.3552 0.01550 0.00560 -0.0950 0.8162 0.4484 0.250 0.3836 0.01541 0.00554 -0.0951 0.8089 0.4555 0.500 0.4111 0.01535 0.00555 -0.0951 0.8012 0.4627 0.750 0.4399 0.01526 0.00556 -0.0952 0.7939 0.4714 1.000 0.4660 0.01522 0.00566 -0.0950 0.7854 0.4816 1.250 0.4951 0.01510 0.00569 -0.0951 0.7785 0.4958 1.500 0.5200 0.01506 0.00589 -0.0946 0.7691 0.5149 1.750 0.5472 0.01491 0.00609 -0.0944 0.7615 0.5484 2.000 0.5682 0.01419 0.00577 -0.0920 0.7303 0.6355 2.500 0.6148 0.01384 0.00481 -0.0873 0.4851 1.0000 2.750 0.6051 0.01727 0.00596 -0.0816 0.0362 1.0000 3.000 0.6269 0.01793 0.00666 -0.0803 0.0242 1.0000 3.250 0.6483 0.01861 0.00753 -0.0789 0.0203 1.0000 3.500 0.6691 0.01932 0.00847 -0.0775 0.0192 1.0000 3.750 0.6881 0.02017 0.00955 -0.0757 0.0186 1.0000 4.000 0.7050 0.02114 0.01070 -0.0736 0.0177 1.0000 4.250 0.7191 0.02224 0.01198 -0.0713 0.0162 1.0000 4.500 0.7304 0.02360 0.01340 -0.0685 0.0150 1.0000 4.750 0.7463 0.02500 0.01480 -0.0663 0.0152 1.0000 5.000 0.7698 0.02654 0.01634 -0.0650 0.0157 1.0000 5.250 0.7997 0.02826 0.01813 -0.0645 0.0168 1.0000 5.500 0.8347 0.03057 0.02050 -0.0647 0.0188 1.0000 5.750 0.8746 0.03413 0.02400 -0.0660 0.0220 1.0000 6.000 0.9035 0.03516 0.02549 -0.0641 0.0281 1.0000