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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4306 0.02590 0.01259 0.0125 1.0000 0.0985 -2.500 -0.3929 0.02402 0.01138 0.0163 1.0000 0.2682 -2.250 -0.3775 0.02332 0.01109 0.0187 1.0000 0.3433 -2.000 -0.3690 0.02191 0.01062 0.0222 1.0000 0.4706 -1.750 -0.0961 0.02382 0.01249 -0.0188 1.0000 1.0000 -1.500 -0.0831 0.02374 0.01221 -0.0166 1.0000 1.0000 -1.250 -0.0704 0.02368 0.01199 -0.0144 1.0000 1.0000 -1.000 -0.0579 0.02363 0.01181 -0.0122 1.0000 1.0000 -0.750 -0.0456 0.02359 0.01167 -0.0099 1.0000 1.0000 -0.500 -0.0334 0.02356 0.01156 -0.0075 1.0000 1.0000 -0.250 -0.0213 0.02353 0.01148 -0.0052 1.0000 1.0000 0.000 -0.0092 0.02352 0.01139 -0.0028 1.0000 1.0000 0.250 0.0028 0.02352 0.01137 -0.0004 1.0000 1.0000 0.500 0.0149 0.02352 0.01139 0.0019 1.0000 1.0000 0.750 0.0270 0.02354 0.01144 0.0043 1.0000 1.0000 1.000 0.0393 0.02356 0.01153 0.0066 1.0000 1.0000 1.250 0.0517 0.02360 0.01168 0.0090 1.0000 1.0000 1.500 0.0643 0.02365 0.01188 0.0113 1.0000 1.0000 1.750 0.0772 0.02372 0.01215 0.0136 1.0000 1.0000 2.250 0.3552 0.02491 0.01020 -0.0239 0.0762 1.0000 2.500 0.3759 0.02537 0.01080 -0.0224 0.0726 1.0000 2.750 0.3969 0.02585 0.01163 -0.0209 0.0705 1.0000 3.000 0.4193 0.02642 0.01241 -0.0197 0.0700 1.0000 3.250 0.4414 0.02702 0.01325 -0.0184 0.0705 1.0000 3.500 0.4628 0.02769 0.01414 -0.0170 0.0717 1.0000 3.750 0.4835 0.02844 0.01511 -0.0153 0.0735 1.0000 4.000 0.5047 0.02931 0.01616 -0.0137 0.0757 1.0000 4.250 0.5292 0.03039 0.01738 -0.0126 0.0784 1.0000 4.500 0.5600 0.03179 0.01885 -0.0123 0.0813 1.0000 4.750 0.5884 0.03238 0.01975 -0.0111 0.0863 1.0000 5.000 0.6219 0.03398 0.02152 -0.0109 0.0923 1.0000