XFOIL Version 6.96 Calculated polar for: NASA SC(2)-0404 AIRFOIL 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 ------ -------- --------- --------- -------- -------- -------- -7.250 -0.6239 0.07154 0.06807 -0.0091 1.0000 0.0093 -7.000 -0.6133 0.06550 0.06194 -0.0144 1.0000 0.0090 -6.750 -0.5993 0.05965 0.05596 -0.0190 1.0000 0.0086 -6.500 -0.5822 0.05398 0.05010 -0.0229 1.0000 0.0083 -6.250 -0.5623 0.04844 0.04432 -0.0261 1.0000 0.0080 -6.000 -0.5396 0.04307 0.03865 -0.0287 1.0000 0.0078 -5.750 -0.5145 0.03790 0.03311 -0.0308 1.0000 0.0076 -5.500 -0.4873 0.03295 0.02773 -0.0323 1.0000 0.0074 -5.250 -0.4583 0.02833 0.02259 -0.0334 1.0000 0.0075 -5.000 -0.4282 0.02420 0.01790 -0.0340 1.0000 0.0078 -4.750 -0.3984 0.02169 0.01480 -0.0341 1.0000 0.0096 -4.500 -0.3701 0.01905 0.01175 -0.0349 1.0000 0.0133 -4.250 -0.3413 0.01683 0.00918 -0.0347 1.0000 0.0138 -4.000 -0.3131 0.01510 0.00721 -0.0345 1.0000 0.0147 -3.750 -0.2851 0.01371 0.00568 -0.0345 1.0000 0.0168 -3.500 -0.2571 0.01292 0.00476 -0.0345 1.0000 0.0206 -3.250 -0.2285 0.01194 0.00369 -0.0350 1.0000 0.0250 -3.000 -0.1999 0.01130 0.00292 -0.0352 1.0000 0.0298 -2.750 -0.1711 0.01067 0.00230 -0.0355 1.0000 0.0549 -2.500 -0.1418 0.00962 0.00181 -0.0367 1.0000 0.2199 -2.250 -0.1153 0.00792 0.00187 -0.0373 1.0000 0.6695 -2.000 -0.0908 0.00778 0.00185 -0.0363 1.0000 0.7293 -1.750 -0.0676 0.00768 0.00186 -0.0349 1.0000 0.7803 -1.500 -0.0458 0.00759 0.00186 -0.0331 1.0000 0.8208 -1.250 -0.0229 0.00751 0.00182 -0.0318 1.0000 0.8439 -1.000 0.0009 0.00743 0.00175 -0.0307 1.0000 0.8609 -0.750 0.0248 0.00737 0.00170 -0.0298 1.0000 0.8769 -0.500 0.0483 0.00729 0.00166 -0.0288 1.0000 0.8922 -0.250 0.0714 0.00721 0.00163 -0.0276 1.0000 0.9097 0.000 0.0945 0.00712 0.00161 -0.0265 1.0000 0.9322 0.250 0.1185 0.00701 0.00160 -0.0257 1.0000 0.9744 0.500 0.1451 0.00703 0.00170 -0.0257 1.0000 1.0000 0.750 0.1738 0.00710 0.00183 -0.0261 1.0000 1.0000 1.000 0.2167 0.00711 0.00196 -0.0295 0.9846 1.0000 1.250 0.2719 0.00773 0.00174 -0.0335 0.6390 1.0000 1.500 0.2892 0.01039 0.00212 -0.0326 0.1060 1.0000 1.750 0.3170 0.01107 0.00253 -0.0329 0.0390 1.0000 2.000 0.3454 0.01167 0.00330 -0.0330 0.0263 1.0000 2.250 0.3734 0.01237 0.00412 -0.0330 0.0232 1.0000 2.500 0.4003 0.01356 0.00541 -0.0328 0.0212 1.0000 2.750 0.4280 0.01427 0.00622 -0.0327 0.0170 1.0000 3.000 0.4554 0.01571 0.00782 -0.0323 0.0146 1.0000 3.250 0.4834 0.01755 0.00991 -0.0318 0.0135 1.0000 3.500 0.5116 0.01984 0.01257 -0.0312 0.0129 1.0000 3.750 0.5371 0.02283 0.01599 -0.0310 0.0092 1.0000 4.000 0.5654 0.02454 0.01809 -0.0304 0.0075 1.0000 4.250 0.5921 0.02862 0.02286 -0.0295 0.0070 1.0000 4.500 0.6174 0.03296 0.02770 -0.0287 0.0069 1.0000 4.750 0.6414 0.03756 0.03274 -0.0282 0.0070 1.0000 5.000 0.6639 0.04239 0.03794 -0.0281 0.0071 1.0000 5.250 0.6850 0.04740 0.04328 -0.0284 0.0073 1.0000 5.500 0.7044 0.05259 0.04874 -0.0292 0.0076 1.0000 5.750 0.7218 0.05793 0.05431 -0.0305 0.0078 1.0000 6.000 0.7368 0.06343 0.06000 -0.0324 0.0081 1.0000 6.250 0.7489 0.06914 0.06586 -0.0348 0.0085 1.0000 6.500 0.7491 0.07611 0.07290 -0.0371 0.0094 1.0000 7.500 0.7957 0.09423 0.09130 -0.0506 0.0104 1.0000 7.750 0.7974 0.09949 0.09655 -0.0559 0.0106 1.0000 8.000 0.7977 0.10449 0.10151 -0.0600 0.0107 1.0000 8.250 0.7983 0.10924 0.10624 -0.0632 0.0108 1.0000