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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1746 0.03379 0.01692 -0.0256 1.0000 1.0000 -2.750 -0.1620 0.03343 0.01615 -0.0239 1.0000 1.0000 -2.500 -0.1486 0.03315 0.01531 -0.0221 1.0000 1.0000 -2.250 -0.1349 0.03292 0.01474 -0.0203 1.0000 1.0000 -2.000 -0.1210 0.03273 0.01422 -0.0185 1.0000 1.0000 -1.750 -0.1070 0.03257 0.01377 -0.0166 1.0000 1.0000 -1.500 -0.0930 0.03243 0.01339 -0.0147 1.0000 1.0000 -1.250 -0.0789 0.03233 0.01307 -0.0128 1.0000 1.0000 -1.000 -0.0649 0.03224 0.01281 -0.0108 1.0000 1.0000 -0.750 -0.0509 0.03217 0.01260 -0.0089 1.0000 1.0000 -0.500 -0.0370 0.03212 0.01245 -0.0069 1.0000 1.0000 -0.250 -0.0230 0.03209 0.01234 -0.0049 1.0000 1.0000 0.000 -0.0091 0.03208 0.01222 -0.0029 1.0000 1.0000 0.250 0.0048 0.03208 0.01220 -0.0009 1.0000 1.0000 0.500 0.0187 0.03210 0.01223 0.0011 1.0000 1.0000 0.750 0.0327 0.03214 0.01232 0.0031 1.0000 1.0000 1.000 0.0466 0.03219 0.01246 0.0051 1.0000 1.0000 1.250 0.0606 0.03227 0.01265 0.0071 1.0000 1.0000 1.500 0.0746 0.03236 0.01291 0.0090 1.0000 1.0000 1.750 0.0886 0.03248 0.01330 0.0110 1.0000 1.0000 2.000 0.1026 0.03263 0.01366 0.0129 1.0000 1.0000 2.250 0.1165 0.03281 0.01409 0.0148 1.0000 1.0000 2.500 0.1303 0.03302 0.01461 0.0166 1.0000 1.0000 2.750 0.1438 0.03328 0.01520 0.0185 1.0000 1.0000 3.000 0.1569 0.03360 0.01589 0.0203 1.0000 1.0000 3.250 0.1693 0.03399 0.01669 0.0222 1.0000 1.0000 3.500 0.1804 0.03451 0.01763 0.0239 1.0000 1.0000 3.750 0.1891 0.03521 0.01877 0.0257 1.0000 1.0000 4.000 0.1925 0.03630 0.02023 0.0273 1.0000 1.0000 4.500 0.4055 0.03438 0.02163 0.0078 0.7317 1.0000 4.750 0.4992 0.03564 0.02069 0.0060 0.2960 1.0000 5.000 0.5164 0.03734 0.02215 0.0085 0.2315 1.0000