XFOIL Version 6.96 Calculated polar for: Eh 1.0/7.0 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.5864 0.09941 0.09310 0.0217 1.0000 0.2571 -7.750 -0.5851 0.09601 0.08976 0.0216 1.0000 0.2722 -7.500 -0.5901 0.09316 0.08702 0.0211 1.0000 0.2883 -7.250 -0.5806 0.08904 0.08294 0.0223 1.0000 0.3070 -7.000 -0.5830 0.08626 0.08025 0.0230 1.0000 0.3291 -6.750 -0.5722 0.08249 0.07643 0.0249 1.0000 0.3550 -6.500 -0.5745 0.07980 0.07383 0.0266 1.0000 0.3843 -6.250 -0.5441 0.07559 0.06959 0.0314 1.0000 0.4338 -5.500 -0.5473 0.05351 0.04642 -0.0115 1.0000 0.1813 -5.250 -0.5182 0.04613 0.03826 -0.0135 1.0000 0.1245 -5.000 -0.4955 0.04150 0.03305 -0.0132 1.0000 0.1100 -4.750 -0.4732 0.03780 0.02887 -0.0126 1.0000 0.1073 -4.500 -0.4497 0.03464 0.02511 -0.0118 1.0000 0.1085 -4.250 -0.4243 0.03157 0.02150 -0.0108 1.0000 0.1084 -4.000 -0.3970 0.02872 0.01816 -0.0098 1.0000 0.1089 -3.750 -0.3684 0.02633 0.01523 -0.0086 1.0000 0.1131 -3.500 -0.3418 0.02404 0.01301 -0.0078 1.0000 0.1317 -3.250 -0.1932 0.01650 0.00856 -0.0188 1.0000 1.0000 -3.000 -0.1742 0.01611 0.00768 -0.0181 1.0000 1.0000 -2.750 -0.1549 0.01578 0.00697 -0.0173 1.0000 1.0000 -2.500 -0.1355 0.01550 0.00638 -0.0164 1.0000 1.0000 -2.250 -0.1159 0.01527 0.00587 -0.0155 1.0000 1.0000 -2.000 -0.0963 0.01507 0.00544 -0.0145 1.0000 1.0000 -1.750 -0.0769 0.01491 0.00502 -0.0134 1.0000 1.0000 -1.500 -0.0576 0.01478 0.00472 -0.0123 1.0000 1.0000 -1.250 -0.0388 0.01469 0.00449 -0.0111 1.0000 1.0000 -1.000 -0.0204 0.01462 0.00431 -0.0098 1.0000 1.0000 -0.750 -0.0028 0.01459 0.00420 -0.0084 1.0000 1.0000 -0.500 0.0137 0.01459 0.00414 -0.0068 1.0000 1.0000 -0.250 0.0287 0.01463 0.00411 -0.0050 1.0000 1.0000 0.000 0.0422 0.01473 0.00418 -0.0030 1.0000 1.0000 0.250 0.0540 0.01491 0.00432 -0.0010 1.0000 1.0000 0.500 0.0642 0.01516 0.00456 0.0012 1.0000 1.0000 0.750 0.0736 0.01552 0.00489 0.0031 1.0000 1.0000 1.000 0.0829 0.01597 0.00531 0.0047 1.0000 1.0000 1.250 0.0927 0.01652 0.00584 0.0059 1.0000 1.0000 1.500 0.1031 0.01715 0.00647 0.0067 1.0000 1.0000 1.750 0.1590 0.01811 0.00754 -0.0011 0.9802 1.0000 2.000 0.2390 0.01890 0.00856 -0.0125 0.9470 1.0000 2.250 0.3270 0.01927 0.00933 -0.0243 0.9150 1.0000 2.500 0.3874 0.01948 0.00983 -0.0300 0.8803 1.0000 2.750 0.4344 0.01964 0.01023 -0.0323 0.8473 1.0000 3.000 0.4631 0.02001 0.01084 -0.0313 0.8137 1.0000 3.250 0.4885 0.02037 0.01134 -0.0294 0.7820 1.0000 3.500 0.5111 0.02074 0.01184 -0.0268 0.7517 1.0000 3.750 0.5312 0.02124 0.01249 -0.0243 0.7199 1.0000 4.000 0.5516 0.02172 0.01312 -0.0216 0.6887 1.0000 4.250 0.5720 0.02214 0.01367 -0.0186 0.6575 1.0000 4.500 0.5918 0.02263 0.01443 -0.0157 0.6228 1.0000 4.750 0.6118 0.02291 0.01485 -0.0123 0.5866 1.0000 5.000 0.6289 0.02262 0.01462 -0.0076 0.5356 1.0000 5.250 0.6365 0.02107 0.01272 -0.0005 0.4364 1.0000 5.500 0.6357 0.02331 0.01310 0.0051 0.1757 1.0000 5.750 0.6540 0.02622 0.01550 0.0071 0.1229 1.0000 6.000 0.6757 0.02857 0.01769 0.0086 0.1007 1.0000 6.250 0.7015 0.03126 0.02039 0.0099 0.0917 1.0000 6.500 0.7278 0.03434 0.02397 0.0112 0.0882 1.0000 6.750 0.7510 0.03760 0.02779 0.0124 0.0857 1.0000 7.000 0.7710 0.04083 0.03141 0.0134 0.0813 1.0000 7.250 0.7895 0.04504 0.03569 0.0140 0.0781 1.0000 7.500 0.8021 0.04888 0.04047 0.0152 0.0803 1.0000 7.750 0.8071 0.05424 0.04659 0.0155 0.0843 1.0000 8.000 0.8111 0.05954 0.05228 0.0155 0.0876 1.0000 8.250 0.8241 0.06483 0.05764 0.0157 0.0911 1.0000 8.500 0.7936 0.07133 0.06482 0.0122 0.0964 1.0000 8.750 0.7849 0.07721 0.07078 0.0100 0.1008 1.0000 9.000 0.7692 0.08331 0.07694 0.0065 0.1061 1.0000 9.250 0.7414 0.09224 0.08577 -0.0026 0.1125 1.0000