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 = 0.070 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4351 0.02301 0.01164 0.0184 1.0000 0.0522 -2.750 -0.4142 0.02218 0.01070 0.0201 1.0000 0.0499 -2.500 -0.3925 0.02150 0.00986 0.0217 1.0000 0.0485 -2.250 -0.3691 0.02073 0.00894 0.0227 1.0000 0.0524 -2.000 -0.3441 0.02014 0.00807 0.0234 1.0000 0.0601 -1.750 -0.3176 0.01950 0.00727 0.0240 1.0000 0.0699 -1.500 -0.2937 0.01863 0.00677 0.0249 1.0000 0.1391 -1.250 -0.2762 0.01767 0.00666 0.0266 1.0000 0.3460 -0.750 -0.0004 0.01774 0.00793 -0.0132 1.0000 1.0000 -0.500 0.0139 0.01774 0.00789 -0.0111 1.0000 1.0000 -0.250 0.0275 0.01777 0.00789 -0.0088 1.0000 1.0000 0.000 0.0405 0.01781 0.00793 -0.0065 1.0000 1.0000 0.250 0.0530 0.01787 0.00801 -0.0040 1.0000 1.0000 0.500 0.0650 0.01795 0.00814 -0.0015 1.0000 1.0000 0.750 0.0766 0.01804 0.00829 0.0011 1.0000 1.0000 1.000 0.0879 0.01814 0.00848 0.0037 1.0000 1.0000 1.250 0.0991 0.01826 0.00871 0.0063 1.0000 1.0000 1.500 0.1101 0.01839 0.00899 0.0090 1.0000 1.0000 1.750 0.1212 0.01854 0.00932 0.0116 1.0000 1.0000 2.250 0.3610 0.02017 0.00776 -0.0196 0.0617 1.0000 2.500 0.3856 0.02058 0.00837 -0.0187 0.0600 1.0000 2.750 0.4100 0.02106 0.00907 -0.0177 0.0597 1.0000 3.000 0.4338 0.02165 0.00987 -0.0167 0.0602 1.0000 3.250 0.4566 0.02238 0.01078 -0.0154 0.0612 1.0000 3.500 0.4780 0.02333 0.01183 -0.0139 0.0624 1.0000 3.750 0.4993 0.02428 0.01309 -0.0124 0.0635 1.0000 4.000 0.5228 0.02489 0.01385 -0.0108 0.0663 1.0000 4.250 0.5480 0.02600 0.01506 -0.0094 0.0705 1.0000 4.500 0.5757 0.02781 0.01688 -0.0085 0.0754 1.0000 4.750 0.6034 0.02883 0.01815 -0.0070 0.0816 1.0000 5.000 0.6320 0.03102 0.02065 -0.0058 0.0910 1.0000