XFOIL Version 6.96 Calculated polar for: WORTMANN FX 2 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -0.7032 0.06904 0.06431 -0.1623 0.8053 0.0065 -19.500 -0.7044 0.06563 0.06081 -0.1642 0.8050 0.0066 -19.250 -0.7050 0.06236 0.05745 -0.1661 0.8047 0.0067 -19.000 -0.7074 0.05897 0.05396 -0.1678 0.8045 0.0067 -18.750 -0.7058 0.05614 0.05105 -0.1692 0.8042 0.0068 -18.500 -0.7054 0.05330 0.04812 -0.1705 0.8039 0.0069 -18.250 -0.7041 0.05065 0.04539 -0.1716 0.8036 0.0070 -18.000 -0.7013 0.04822 0.04289 -0.1726 0.8033 0.0071 -17.750 -0.6957 0.04616 0.04076 -0.1734 0.8030 0.0072 -17.500 -0.6896 0.04420 0.03875 -0.1741 0.8026 0.0074 -17.250 -0.6816 0.04247 0.03696 -0.1747 0.8023 0.0075 -17.000 -0.6719 0.04092 0.03536 -0.1752 0.8020 0.0077 -16.750 -0.6627 0.03937 0.03376 -0.1756 0.8017 0.0079 -16.500 -0.6520 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0.0124 0.6976 13.250 1.1617 0.04063 0.03349 -0.0331 0.0117 0.6982 13.500 1.1746 0.04173 0.03462 -0.0322 0.0111 0.6988 13.750 1.1872 0.04285 0.03578 -0.0313 0.0106 0.6993 14.000 1.2011 0.04389 0.03686 -0.0306 0.0101 0.6998 14.250 1.2138 0.04504 0.03806 -0.0299 0.0098 0.7003 14.500 1.2265 0.04622 0.03928 -0.0291 0.0093 0.7008 14.750 1.2385 0.04745 0.04055 -0.0284 0.0090 0.7013 15.000 1.2503 0.04873 0.04188 -0.0277 0.0087 0.7018 15.250 1.2608 0.05013 0.04332 -0.0270 0.0083 0.7023 15.500 1.2705 0.05164 0.04488 -0.0262 0.0080 0.7030 15.750 1.2815 0.05306 0.04636 -0.0256 0.0077 0.7035 16.000 1.2918 0.05456 0.04792 -0.0250 0.0075 0.7040 16.250 1.3024 0.05604 0.04945 -0.0245 0.0073 0.7045 16.500 1.3106 0.05779 0.05127 -0.0239 0.0072 0.7048 16.750 1.3198 0.05946 0.05299 -0.0234 0.0069 0.7051 17.000 1.3281 0.06124 0.05483 -0.0229 0.0067 0.7054 17.250 1.3352 0.06318 0.05682 -0.0225 0.0065 0.7058 17.500 1.3421 0.06517 0.05887 -0.0221 0.0063 0.7061 17.750 1.3477 0.06735 0.06112 -0.0217 0.0062 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