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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1731 0.02506 0.01239 -0.0249 1.0000 1.0000 -2.750 -0.1541 0.02483 0.01171 -0.0235 1.0000 1.0000 -2.500 -0.1355 0.02464 0.01111 -0.0221 1.0000 1.0000 -2.250 -0.1172 0.02448 0.01057 -0.0206 1.0000 1.0000 -2.000 -0.0994 0.02435 0.01012 -0.0191 1.0000 1.0000 -1.750 -0.0821 0.02423 0.00973 -0.0175 1.0000 1.0000 -1.500 -0.0651 0.02414 0.00943 -0.0159 1.0000 1.0000 -1.250 -0.0486 0.02407 0.00919 -0.0142 1.0000 1.0000 -1.000 -0.0324 0.02402 0.00891 -0.0124 1.0000 1.0000 -0.750 -0.0164 0.02399 0.00876 -0.0106 1.0000 1.0000 -0.500 -0.0008 0.02398 0.00867 -0.0087 1.0000 1.0000 -0.250 0.0145 0.02399 0.00864 -0.0068 1.0000 1.0000 0.000 0.0295 0.02401 0.00865 -0.0048 1.0000 1.0000 0.250 0.0441 0.02406 0.00871 -0.0027 1.0000 1.0000 0.500 0.0585 0.02412 0.00881 -0.0007 1.0000 1.0000 0.750 0.0726 0.02419 0.00899 0.0015 1.0000 1.0000 1.000 0.0865 0.02429 0.00918 0.0036 1.0000 1.0000 1.250 0.1002 0.02440 0.00942 0.0058 1.0000 1.0000 1.500 0.1137 0.02454 0.00972 0.0079 1.0000 1.0000 1.750 0.1271 0.02469 0.01007 0.0101 1.0000 1.0000 2.000 0.1403 0.02487 0.01048 0.0122 1.0000 1.0000 2.250 0.1534 0.02508 0.01096 0.0144 1.0000 1.0000 2.500 0.1664 0.02532 0.01160 0.0165 1.0000 1.0000 2.750 0.1793 0.02559 0.01222 0.0186 1.0000 1.0000 3.000 0.1919 0.02590 0.01291 0.0207 1.0000 1.0000 3.250 0.2039 0.02628 0.01369 0.0228 1.0000 1.0000 3.750 0.4502 0.02749 0.01299 -0.0027 0.1499 1.0000 4.000 0.4673 0.02894 0.01438 -0.0006 0.1280 1.0000 4.250 0.4859 0.03017 0.01562 0.0013 0.1215 1.0000 4.500 0.5054 0.03133 0.01696 0.0033 0.1182 1.0000 4.750 0.5262 0.03262 0.01845 0.0053 0.1165 1.0000 5.000 0.5502 0.03407 0.02014 0.0069 0.1161 1.0000