XFOIL Version 6.94 Calculated polar for: EPPLER 864 STRUT 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 0.0025 0.02112 0.01324 -0.0004 0.4363 0.4373 0.500 0.0553 0.02122 0.01338 -0.0008 0.4344 0.4402 1.000 0.1074 0.02116 0.01334 -0.0010 0.4320 0.4430 1.500 0.1591 0.02114 0.01332 -0.0012 0.4290 0.4452 2.000 0.2117 0.02118 0.01334 -0.0015 0.4259 0.4468 2.500 0.2647 0.02083 0.01298 -0.0020 0.4230 0.4506 3.000 0.3201 0.02066 0.01282 -0.0028 0.4200 0.4541 3.500 0.3758 0.02086 0.01300 -0.0038 0.4166 0.4569 4.000 0.4251 0.02138 0.01358 -0.0038 0.4137 0.4598 4.500 0.4692 0.02151 0.01380 -0.0027 0.4116 0.4628 5.000 0.5102 0.02168 0.01403 -0.0012 0.4087 0.4659 5.500 0.5509 0.02187 0.01425 0.0004 0.4057 0.4687 6.000 0.5874 0.02203 0.01440 0.0028 0.4030 0.4706 6.500 0.6287 0.02197 0.01434 0.0042 0.4004 0.4745 7.000 0.6788 0.02211 0.01449 0.0040 0.3974 0.4790 7.500 0.7150 0.02293 0.01540 0.0056 0.3942 0.4827 8.000 0.7297 0.02393 0.01656 0.0099 0.3912 0.4863 8.500 0.7528 0.02497 0.01770 0.0125 0.3875 0.4903 9.000 0.7899 0.02566 0.01840 0.0132 0.3839 0.4940 9.500 0.8374 0.02606 0.01874 0.0127 0.3806 0.4968 10.000 0.8832 0.02672 0.01944 0.0122 0.3769 0.5040 10.500 0.8635 0.03023 0.02324 0.0168 0.3723 0.5085 11.000 0.8773 0.03268 0.02581 0.0179 0.3675 0.5137 11.500 0.9192 0.03367 0.02680 0.0171 0.3638 0.5196 12.000 0.9552 0.03521 0.02832 0.0167 0.3599 0.5246 12.500 0.8800 0.04482 0.03826 0.0204 0.3505 0.5290 13.000 0.9330 0.04514 0.03862 0.0187 0.3470 0.5400 14.000 0.8495 0.06220 0.05599 0.0207 0.3267 0.5507