XFOIL Version 6.94 Calculated polar for: voyager2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.080 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4758 0.00504 -0.00476 0.0208 1.0000 0.0664 -2.750 -0.4564 0.00415 -0.00561 0.0226 1.0000 0.0654 -2.500 -0.4375 0.00351 -0.00631 0.0244 1.0000 0.0636 -2.250 -0.4173 0.00307 -0.00688 0.0259 1.0000 0.0630 -2.000 -0.3955 0.00272 -0.00739 0.0271 1.0000 0.0642 -1.750 -0.3732 0.00246 -0.00780 0.0282 1.0000 0.0674 -1.500 -0.3507 0.00216 -0.00804 0.0293 1.0000 0.0810 -1.250 -0.3252 0.01917 0.00854 0.0302 1.0000 0.3161 -0.750 -0.0490 0.02012 0.01148 -0.0090 1.0000 0.9987 -0.500 -0.0311 0.02011 0.01138 -0.0079 1.0000 1.0000 -0.250 -0.0201 0.02010 0.01133 -0.0053 1.0000 1.0000 0.000 -0.0092 0.02009 0.01125 -0.0028 1.0000 1.0000 0.250 0.0016 0.02008 0.01123 -0.0002 1.0000 1.0000 0.500 0.0125 0.02007 0.01123 0.0023 1.0000 1.0000 0.750 0.0235 0.02007 0.01126 0.0049 1.0000 1.0000 1.000 0.0346 0.02007 0.01134 0.0074 1.0000 1.0000 1.250 0.3160 0.02041 0.00806 -0.0372 0.0454 1.0000 1.500 0.3393 0.02060 0.00839 -0.0363 0.0453 1.0000 1.750 0.3629 0.02086 0.00881 -0.0353 0.0461 1.0000 2.000 0.3863 0.02120 0.00933 -0.0344 0.0474 1.0000 2.250 0.4092 0.02165 0.00996 -0.0333 0.0491 1.0000 2.500 0.4308 0.02221 0.01080 -0.0321 0.0504 1.0000 2.750 0.4514 0.02274 0.01147 -0.0306 0.0523 1.0000 3.000 0.4725 0.02308 0.01196 -0.0290 0.0558 1.0000 3.250 0.4924 0.02375 0.01274 -0.0272 0.0598 1.0000 3.500 0.5130 0.02485 0.01379 -0.0255 0.0639 1.0000 3.750 0.5368 0.02557 0.01453 -0.0242 0.0685 1.0000 4.000 0.5636 0.02639 0.01547 -0.0230 0.0763 1.0000 4.250 0.5925 0.02747 0.01656 -0.0222 0.0838 1.0000 4.500 0.6249 0.02933 0.01851 -0.0219 0.0952 1.0000 4.750 0.6524 0.02984 0.01945 -0.0201 0.1086 1.0000 5.000 0.6830 0.03130 0.02125 -0.0188 0.1249 1.0000