XFOIL Version 6.96 Calculated polar for: RAF 89 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.9535 0.08710 0.08087 -0.0813 1.0000 0.1071 -19.500 -0.9825 0.08067 0.07419 -0.0839 1.0000 0.1085 -19.250 -0.9919 0.07683 0.07022 -0.0850 1.0000 0.1098 -19.000 -0.9670 0.07719 0.07072 -0.0842 1.0000 0.1116 -18.750 -0.9594 0.07546 0.06901 -0.0844 1.0000 0.1132 -18.500 -0.9611 0.07266 0.06615 -0.0850 1.0000 0.1151 -18.250 -0.9714 0.06892 0.06227 -0.0859 1.0000 0.1171 -18.000 -0.9850 0.06496 0.05811 -0.0868 1.0000 0.1187 -17.750 -0.9633 0.06493 0.05826 -0.0860 1.0000 0.1211 -17.500 -0.9579 0.06316 0.05652 -0.0860 1.0000 0.1235 -17.250 -0.9593 0.06070 0.05400 -0.0861 1.0000 0.1259 -17.000 -0.9693 0.05747 0.05060 -0.0863 1.0000 0.1284 -16.750 -0.9564 0.05661 0.04987 -0.0856 1.0000 0.1313 -16.500 -0.9526 0.05491 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1.1507 0.02980 0.02277 0.0113 0.3199 1.0000 12.500 1.1698 0.03020 0.02306 0.0122 0.3144 1.0000 12.750 1.1589 0.03164 0.02463 0.0157 0.3092 1.0000 13.000 1.1626 0.03257 0.02557 0.0178 0.3036 1.0000 13.250 1.1876 0.03274 0.02558 0.0180 0.2984 1.0000 13.500 1.1782 0.03444 0.02743 0.0208 0.2938 1.0000 13.750 1.1792 0.03574 0.02879 0.0227 0.2887 1.0000 14.000 1.1914 0.03649 0.02949 0.0237 0.2840 1.0000 14.250 1.2047 0.03731 0.03027 0.0246 0.2794 1.0000 14.500 1.1983 0.03922 0.03233 0.0266 0.2747 1.0000 14.750 1.2035 0.04049 0.03361 0.0278 0.2698 1.0000 15.000 1.2269 0.04069 0.03366 0.0280 0.2652 1.0000 15.250 1.2168 0.04307 0.03621 0.0298 0.2606 1.0000 15.500 1.2155 0.04496 0.03817 0.0309 0.2556 1.0000 15.750 1.2273 0.04591 0.03908 0.0316 0.2511 1.0000 16.000 1.2337 0.04737 0.04054 0.0323 0.2464 1.0000 16.250 1.2249 0.05004 0.04334 0.0334 0.2414 1.0000 16.500 1.2311 0.05153 0.04482 0.0339 0.2365 1.0000 16.750 1.2435 0.05257 0.04581 0.0344 0.2319 1.0000 17.000 1.2294 0.05599 0.04940 0.0351 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