XFOIL Version 6.94 Calculated polar for: GOE 484 AIRFOIL 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 0.5891 0.01069 0.00405 -0.1421 0.8858 0.0529 0.500 0.6489 0.01014 0.00352 -0.1422 0.8708 0.0613 1.000 0.7048 0.00943 0.00305 -0.1411 0.8419 0.1576 1.500 0.7601 0.00763 0.00297 -0.1404 0.8057 1.0000 2.000 0.8099 0.00772 0.00285 -0.1383 0.7581 1.0000 2.500 0.8544 0.00809 0.00275 -0.1350 0.6664 1.0000 3.000 0.8901 0.00924 0.00312 -0.1304 0.5427 1.0000 3.500 0.8925 0.01353 0.00484 -0.1212 0.0622 1.0000 4.000 0.9389 0.01434 0.00553 -0.1194 0.0442 1.0000 4.500 0.9855 0.01508 0.00622 -0.1177 0.0359 1.0000 5.000 1.0326 0.01571 0.00686 -0.1160 0.0317 1.0000 5.500 1.0803 0.01620 0.00730 -0.1146 0.0231 1.0000 6.000 1.1257 0.01691 0.00795 -0.1127 0.0177 1.0000 6.500 1.1681 0.01787 0.00891 -0.1103 0.0063 1.0000 7.000 1.2074 0.01905 0.01022 -0.1073 0.0057 1.0000 7.500 1.2459 0.02020 0.01155 -0.1042 0.0053 1.0000 8.000 1.2816 0.02142 0.01305 -0.1005 0.0052 1.0000 8.500 1.3133 0.02287 0.01472 -0.0963 0.0052 1.0000 9.000 1.3423 0.02459 0.01666 -0.0918 0.0052 1.0000 9.500 1.3679 0.02670 0.01906 -0.0871 0.0053 1.0000 10.000 1.3893 0.02934 0.02205 -0.0821 0.0054 1.0000 10.500 1.4066 0.03267 0.02579 -0.0770 0.0055 1.0000 11.000 1.4183 0.03723 0.03096 -0.0717 0.0057 1.0000 11.500 1.4204 0.04331 0.03770 -0.0664 0.0059 1.0000 12.000 1.4044 0.05128 0.04640 -0.0610 0.0060 1.0000 12.500 1.3727 0.06090 0.05667 -0.0574 0.0061 1.0000 13.000 1.3304 0.07268 0.06903 -0.0572 0.0062 1.0000 13.500 1.2835 0.08715 0.08397 -0.0618 0.0062 1.0000 14.000 1.2401 0.10399 0.10119 -0.0709 0.0062 1.0000 15.000 1.1262 0.16445 0.16190 -0.1074 0.0063 1.0000 15.500 1.1140 0.18082 0.17823 -0.1166 0.0065 1.0000 16.000 1.1168 0.19247 0.18985 -0.1230 0.0069 1.0000 16.500 1.1114 0.21062 0.20791 -0.1322 0.0088 1.0000 17.000 1.1197 0.22099 0.21825 -0.1379 0.0103 1.0000 19.000 0.9209 0.25500 0.25282 -0.1355 0.0190 1.0000 19.500 0.9247 0.26371 0.26155 -0.1390 0.0189 1.0000 20.000 0.9265 0.27190 0.26975 -0.1432 0.0187 1.0000 20.500 0.9291 0.28014 0.27800 -0.1476 0.0181 1.0000 21.000 0.9343 0.28861 0.28648 -0.1511 0.0166 1.0000 21.500 0.9399 0.29690 0.29479 -0.1542 0.0152 1.0000 22.000 0.9456 0.30488 0.30284 -0.1571 0.0142 1.0000 22.500 0.9514 0.31277 0.31075 -0.1599 0.0134 1.0000 23.000 0.9575 0.32068 0.31869 -0.1624 0.0127 1.0000 23.500 0.9672 0.33186 0.32989 -0.1632 0.0119 1.0000 24.000 0.9710 0.34046 0.33851 -0.1664 0.0119 1.0000 24.500 0.9741 0.34886 0.34693 -0.1698 0.0118 1.0000 25.000 0.9763 0.35757 0.35565 -0.1734 0.0118 1.0000 25.500 0.9797 0.36596 0.36407 -0.1766 0.0117 1.0000 26.000 0.9830 0.37446 0.37262 -0.1801 0.0115 1.0000 26.500 0.9872 0.38349 0.38168 -0.1834 0.0109 1.0000 27.000 0.9912 0.39237 0.39059 -0.1862 0.0097 1.0000 27.500 0.9945 0.40116 0.39940 -0.1890 0.0095 1.0000 28.000 0.9977 0.40978 0.40806 -0.1918 0.0089 1.0000