XFOIL Version 6.94 Calculated polar for: AIRFOIL 3024 9.84% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4236 0.00802 0.00330 -0.0868 0.7949 1.0000 0.500 0.4727 0.00801 0.00312 -0.0851 0.7782 1.0000 1.000 0.5225 0.00803 0.00300 -0.0835 0.7602 1.0000 1.500 0.5727 0.00807 0.00293 -0.0821 0.7407 1.0000 2.000 0.6233 0.00814 0.00288 -0.0807 0.7197 1.0000 2.500 0.6743 0.00825 0.00290 -0.0795 0.6963 1.0000 3.000 0.7249 0.00840 0.00297 -0.0782 0.6683 1.0000 3.500 0.7753 0.00861 0.00311 -0.0769 0.6369 1.0000 4.000 0.8251 0.00891 0.00329 -0.0756 0.6013 1.0000 4.500 0.8738 0.00932 0.00359 -0.0741 0.5612 1.0000 5.000 0.9212 0.00982 0.00397 -0.0725 0.5161 1.0000 5.500 0.9673 0.01044 0.00447 -0.0708 0.4681 1.0000 6.000 1.0114 0.01118 0.00505 -0.0688 0.4161 1.0000 6.500 1.0528 0.01209 0.00577 -0.0665 0.3609 1.0000 7.000 1.0905 0.01320 0.00660 -0.0638 0.2945 1.0000 7.500 1.1285 0.01431 0.00753 -0.0612 0.2417 1.0000 8.000 1.1646 0.01550 0.00858 -0.0584 0.1946 1.0000 8.500 1.1974 0.01684 0.00975 -0.0552 0.1474 1.0000 9.000 1.2229 0.01843 0.01114 -0.0509 0.0977 1.0000 9.500 1.2300 0.02118 0.01342 -0.0444 0.0335 1.0000 10.000 1.2424 0.02383 0.01610 -0.0390 0.0220 1.0000 10.500 1.2598 0.02621 0.01871 -0.0349 0.0195 1.0000 11.000 1.2706 0.02921 0.02193 -0.0307 0.0179 1.0000 11.500 1.2735 0.03310 0.02602 -0.0268 0.0171 1.0000 12.000 1.2716 0.03789 0.03102 -0.0232 0.0165 1.0000 12.500 1.2819 0.04185 0.03523 -0.0210 0.0157 1.0000 13.000 1.2887 0.04649 0.04012 -0.0192 0.0151 1.0000 13.500 1.2944 0.05164 0.04557 -0.0177 0.0148 1.0000 14.000 1.2972 0.05750 0.05176 -0.0166 0.0146 1.0000 14.500 1.2949 0.06432 0.05894 -0.0165 0.0145 1.0000 15.000 1.2846 0.07262 0.06764 -0.0176 0.0144 1.0000 15.500 1.2650 0.08289 0.07836 -0.0206 0.0145 1.0000 16.000 1.2368 0.09551 0.09141 -0.0259 0.0146 1.0000 16.500 1.2028 0.11037 0.10669 -0.0337 0.0149 1.0000 17.000 1.1639 0.12790 0.12460 -0.0443 0.0151 1.0000 17.500 1.1211 0.14833 0.14536 -0.0573 0.0155 1.0000 18.000 1.0693 0.17424 0.17151 -0.0736 0.0161 1.0000