XFOIL Version 6.94 Calculated polar for: Voyager1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1841 0.03631 0.01493 -0.0212 1.0000 1.0000 -2.750 -0.1667 0.03601 0.01408 -0.0199 1.0000 1.0000 -2.500 -0.1492 0.03575 0.01332 -0.0186 1.0000 1.0000 -2.250 -0.1318 0.03553 0.01266 -0.0172 1.0000 1.0000 -2.000 -0.1144 0.03535 0.01209 -0.0159 1.0000 1.0000 -1.750 -0.0971 0.03520 0.01160 -0.0144 1.0000 1.0000 -1.500 -0.0800 0.03507 0.01119 -0.0130 1.0000 1.0000 -1.250 -0.0630 0.03497 0.01086 -0.0114 1.0000 1.0000 -1.000 -0.0462 0.03490 0.01046 -0.0099 1.0000 1.0000 -0.750 -0.0295 0.03485 0.01025 -0.0083 1.0000 1.0000 -0.500 -0.0129 0.03482 0.01010 -0.0067 1.0000 1.0000 -0.250 0.0035 0.03481 0.01002 -0.0051 1.0000 1.0000 0.000 0.0199 0.03483 0.01000 -0.0034 1.0000 1.0000 0.250 0.0361 0.03486 0.01005 -0.0018 1.0000 1.0000 0.500 0.0522 0.03492 0.01016 -0.0001 1.0000 1.0000 0.750 0.0681 0.03500 0.01037 0.0016 1.0000 1.0000 1.000 0.0838 0.03510 0.01060 0.0034 1.0000 1.0000 1.250 0.0994 0.03523 0.01089 0.0051 1.0000 1.0000 1.500 0.1147 0.03539 0.01125 0.0068 1.0000 1.0000 1.750 0.1299 0.03557 0.01168 0.0086 1.0000 1.0000 2.000 0.1448 0.03578 0.01219 0.0103 1.0000 1.0000 2.250 0.1594 0.03602 0.01276 0.0121 1.0000 1.0000 2.500 0.1737 0.03631 0.01356 0.0138 1.0000 1.0000 2.750 0.1876 0.03663 0.01428 0.0155 1.0000 1.0000 3.000 0.2010 0.03701 0.01509 0.0173 1.0000 1.0000 3.250 0.2139 0.03745 0.01600 0.0190 1.0000 1.0000 3.500 0.2259 0.03797 0.01701 0.0207 1.0000 1.0000 3.750 0.2369 0.03859 0.01816 0.0223 1.0000 1.0000 4.000 0.2464 0.03935 0.01944 0.0239 1.0000 1.0000 4.250 0.2538 0.04030 0.02089 0.0254 1.0000 1.0000 4.500 0.2584 0.04150 0.02254 0.0267 1.0000 1.0000 4.750 0.2604 0.04299 0.02438 0.0278 1.0000 1.0000 5.000 0.2617 0.04467 0.02634 0.0285 1.0000 1.0000