XFOIL Version 6.94 Calculated polar for: EPPLER 664 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3596 0.01518 0.01041 -0.0703 0.7393 0.8182 0.500 0.4098 0.01518 0.01039 -0.0696 0.7334 0.8221 1.000 0.4551 0.01494 0.01016 -0.0684 0.7269 0.8272 1.500 0.5110 0.01461 0.00976 -0.0698 0.7208 0.8316 2.000 0.5691 0.01437 0.00944 -0.0711 0.7148 0.8343 2.500 0.6130 0.01401 0.00917 -0.0692 0.7072 0.8362 3.000 0.6676 0.01369 0.00885 -0.0693 0.6995 0.8381 3.500 0.7219 0.01348 0.00864 -0.0695 0.6915 0.8401 4.000 0.7712 0.01314 0.00837 -0.0688 0.6812 0.8421 4.500 0.8268 0.01282 0.00806 -0.0694 0.6710 0.8440 5.000 0.8745 0.01249 0.00781 -0.0684 0.6573 0.8466 5.500 0.9212 0.01224 0.00764 -0.0672 0.6408 0.8490 6.000 0.9659 0.01203 0.00747 -0.0657 0.6187 0.8509 6.500 1.0022 0.01198 0.00740 -0.0625 0.5876 0.8527 7.000 1.0174 0.01202 0.00732 -0.0551 0.5461 0.8551 7.500 1.0245 0.01273 0.00786 -0.0464 0.4965 0.8582 8.000 1.0257 0.01390 0.00882 -0.0377 0.4422 0.8612 8.500 1.0294 0.01537 0.01012 -0.0303 0.3934 0.8639 9.000 1.0314 0.01727 0.01180 -0.0236 0.3412 0.8666 9.500 1.0329 0.01957 0.01383 -0.0175 0.2850 0.8691 10.000 1.0450 0.02168 0.01576 -0.0134 0.2386 0.8711 10.500 1.0598 0.02380 0.01774 -0.0098 0.1996 0.8733 11.000 1.0717 0.02608 0.01984 -0.0062 0.1601 0.8755 11.500 1.0876 0.02833 0.02199 -0.0032 0.1290 0.8775 12.000 1.1045 0.03069 0.02429 -0.0006 0.1031 0.8796 12.500 1.1186 0.03340 0.02691 0.0018 0.0781 0.8817 13.000 1.1351 0.03613 0.02962 0.0038 0.0593 0.8839 13.500 1.1504 0.03913 0.03261 0.0054 0.0444 0.8861 14.000 1.1636 0.04250 0.03598 0.0069 0.0310 0.8883 14.500 1.1727 0.04646 0.03996 0.0082 0.0197 0.8905 15.500 1.1816 0.05605 0.04978 0.0097 0.0105 0.8952 16.000 1.1860 0.06127 0.05522 0.0098 0.0091 0.8977 16.500 1.1862 0.06734 0.06148 0.0091 0.0085 0.9002 17.000 1.1813 0.07448 0.06885 0.0078 0.0081 0.9026 17.500 1.1770 0.08199 0.07659 0.0057 0.0080 0.9050 18.000 1.1706 0.09019 0.08504 0.0028 0.0078 0.9073 18.500 1.1618 0.09918 0.09427 -0.0008 0.0076 0.9095 19.500 1.1458 0.11748 0.11302 -0.0094 0.0074 0.9149 20.000 1.1387 0.12643 0.12219 -0.0139 0.0073 0.9188 20.500 1.1336 0.13516 0.13113 -0.0186 0.0073 0.9237 21.000 1.1292 0.14373 0.13991 -0.0235 0.0072 0.9305 22.000 1.1219 0.16071 0.15736 -0.0345 0.0072 1.0000 22.500 1.1157 0.17026 0.16714 -0.0410 0.0072 1.0000 23.000 1.1056 0.18077 0.17788 -0.0483 0.0073 1.0000 23.500 1.0876 0.19327 0.19065 -0.0568 0.0074 1.0000 24.000 1.0533 0.21043 0.20813 -0.0678 0.0076 1.0000