XFOIL Version 6.94 Calculated polar for: S6061 9% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.1681 0.00596 0.00184 -0.0315 0.7738 0.9349 0.500 0.2317 0.00600 0.00178 -0.0326 0.7489 0.9665 1.000 0.3108 0.00609 0.00172 -0.0374 0.7225 0.9848 1.500 0.3930 0.00616 0.00168 -0.0431 0.6933 0.9992 2.000 0.4403 0.00628 0.00167 -0.0416 0.6620 1.0000 2.500 0.4873 0.00646 0.00176 -0.0399 0.6294 1.0000 3.000 0.5363 0.00670 0.00190 -0.0386 0.5942 1.0000 3.500 0.5866 0.00699 0.00209 -0.0375 0.5521 1.0000 4.000 0.6360 0.00748 0.00234 -0.0363 0.4801 1.0000 4.500 0.6844 0.00822 0.00271 -0.0352 0.3864 1.0000 5.000 0.7326 0.00911 0.00326 -0.0342 0.2944 1.0000 5.500 0.7795 0.01025 0.00395 -0.0332 0.1930 1.0000 6.000 0.8238 0.01178 0.00493 -0.0319 0.0827 1.0000 6.500 0.8664 0.01363 0.00652 -0.0300 0.0238 1.0000 7.000 0.9126 0.01490 0.00799 -0.0285 0.0212 1.0000 7.500 0.9561 0.01638 0.00964 -0.0267 0.0191 1.0000 8.000 0.9946 0.01843 0.01188 -0.0242 0.0179 1.0000 8.500 1.0280 0.02150 0.01520 -0.0211 0.0169 1.0000 9.000 1.0628 0.02527 0.01935 -0.0185 0.0155 1.0000 9.500 1.0980 0.02857 0.02303 -0.0160 0.0146 1.0000 10.000 1.1233 0.03320 0.02818 -0.0126 0.0137 1.0000 10.500 1.1408 0.03664 0.03196 -0.0089 0.0121 1.0000 11.000 1.1414 0.03967 0.03522 -0.0038 0.0109 1.0000 11.500 1.1233 0.04549 0.04145 0.0007 0.0105 1.0000 12.000 1.0868 0.05417 0.05057 0.0022 0.0101 1.0000 12.500 1.0592 0.06297 0.05974 0.0000 0.0102 1.0000 13.000 1.0278 0.07433 0.07143 -0.0060 0.0103 1.0000 13.500 0.9905 0.08943 0.08681 -0.0160 0.0103 1.0000