XFOIL Version 6.96 Calculated polar for: NACA 65-206 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.750 -0.5296 0.08236 0.07887 -0.0190 1.0000 0.0103 -7.500 -0.5268 0.07738 0.07389 -0.0229 1.0000 0.0103 -7.250 -0.5332 0.07170 0.06823 -0.0247 1.0000 0.0092 -7.000 -0.5284 0.06700 0.06348 -0.0276 1.0000 0.0089 -6.750 -0.5210 0.06237 0.05878 -0.0303 1.0000 0.0085 -6.500 -0.5112 0.05774 0.05404 -0.0325 1.0000 0.0082 -6.250 -0.4992 0.05304 0.04919 -0.0343 1.0000 0.0080 -6.000 -0.4851 0.04852 0.04448 -0.0354 1.0000 0.0079 -5.750 -0.4691 0.04409 0.03982 -0.0360 1.0000 0.0079 -5.500 -0.4504 0.04017 0.03564 -0.0360 1.0000 0.0085 -5.000 -0.4093 0.03274 0.02750 -0.0347 1.0000 0.0109 -4.750 -0.3883 0.02991 0.02435 -0.0338 1.0000 0.0107 -4.500 -0.3673 0.02712 0.02122 -0.0329 1.0000 0.0106 -4.250 -0.3460 0.02422 0.01793 -0.0319 1.0000 0.0106 -4.000 -0.3239 0.02134 0.01462 -0.0309 1.0000 0.0107 -3.750 -0.3012 0.01840 0.01125 -0.0300 1.0000 0.0114 -3.500 -0.2784 0.01668 0.00930 -0.0290 1.0000 0.0126 -3.250 -0.2480 0.01596 0.00843 -0.0299 0.9971 0.0172 -3.000 -0.2149 0.01439 0.00659 -0.0307 0.9934 0.0194 -2.750 -0.1819 0.01338 0.00546 -0.0319 0.9892 0.0225 -2.500 -0.1490 0.01233 0.00437 -0.0333 0.9850 0.0268 -2.250 -0.1153 0.01164 0.00357 -0.0348 0.9802 0.0301 -2.000 -0.0809 0.01116 0.00294 -0.0363 0.9757 0.0361 -1.750 -0.0474 0.01060 0.00244 -0.0378 0.9702 0.0707 -1.500 -0.0260 0.00808 0.00243 -0.0377 0.9653 0.7271 -1.250 -0.0035 0.00788 0.00252 -0.0356 0.9576 0.8360 -1.000 0.0172 0.00780 0.00256 -0.0330 0.9497 0.9058 -0.750 0.0487 0.00775 0.00246 -0.0334 0.9437 0.9339 -0.500 0.0834 0.00771 0.00233 -0.0349 0.9374 0.9471 -0.250 0.1198 0.00766 0.00223 -0.0368 0.9312 0.9592 0.000 0.1574 0.00763 0.00215 -0.0389 0.9248 0.9709 0.250 0.1954 0.00759 0.00210 -0.0413 0.9180 0.9829 0.500 0.2323 0.00756 0.00209 -0.0434 0.9101 0.9990 0.750 0.2614 0.00757 0.00209 -0.0437 0.9006 1.0000 1.000 0.2887 0.00760 0.00213 -0.0437 0.8901 1.0000 1.250 0.3155 0.00764 0.00221 -0.0436 0.8789 1.0000 1.500 0.3422 0.00768 0.00229 -0.0433 0.8661 1.0000 1.750 0.3664 0.00768 0.00229 -0.0419 0.8349 1.0000 2.000 0.3890 0.00773 0.00224 -0.0402 0.7854 1.0000 2.250 0.4062 0.00817 0.00205 -0.0370 0.6380 1.0000 2.500 0.4147 0.01039 0.00237 -0.0340 0.2190 1.0000 2.750 0.4352 0.01176 0.00297 -0.0333 0.0467 1.0000 3.000 0.4607 0.01232 0.00364 -0.0329 0.0356 1.0000 3.250 0.4860 0.01289 0.00435 -0.0325 0.0318 1.0000 3.500 0.5100 0.01373 0.00526 -0.0320 0.0270 1.0000 3.750 0.5345 0.01448 0.00610 -0.0315 0.0242 1.0000 4.000 0.5587 0.01548 0.00721 -0.0308 0.0218 1.0000 4.250 0.5830 0.01682 0.00867 -0.0300 0.0198 1.0000 4.500 0.6077 0.01776 0.00965 -0.0297 0.0152 1.0000 4.750 0.6319 0.02045 0.01259 -0.0288 0.0135 1.0000 5.000 0.6580 0.02244 0.01500 -0.0279 0.0126 1.0000 5.250 0.6837 0.02449 0.01744 -0.0270 0.0106 1.0000 5.500 0.7081 0.02625 0.01950 -0.0263 0.0084 1.0000 5.750 0.7305 0.02917 0.02285 -0.0252 0.0078 1.0000 6.000 0.7509 0.03267 0.02680 -0.0239 0.0075 1.0000 6.250 0.7690 0.03674 0.03132 -0.0225 0.0075 1.0000 6.500 0.7848 0.04158 0.03662 -0.0209 0.0076 1.0000 6.750 0.7979 0.04676 0.04222 -0.0196 0.0079 1.0000 7.000 0.8083 0.05206 0.04787 -0.0186 0.0083 1.0000 7.250 0.8156 0.05741 0.05352 -0.0180 0.0088 1.0000 7.500 0.8200 0.06260 0.05895 -0.0179 0.0091 1.0000 7.750 0.8212 0.06771 0.06425 -0.0182 0.0095 1.0000 8.000 0.8185 0.07287 0.06955 -0.0192 0.0099 1.0000 8.250 0.8129 0.07778 0.07456 -0.0206 0.0101 1.0000 8.500 0.8017 0.08249 0.07932 -0.0224 0.0103 1.0000 8.750 0.7916 0.08817 0.08502 -0.0271 0.0104 1.0000