XFOIL Version 6.96 Calculated polar for: NASA SC(2)-0406 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 ------ -------- --------- --------- -------- -------- -------- -9.000 -0.6885 0.08694 0.08346 0.0052 1.0000 0.0137 -8.750 -0.6936 0.08032 0.07688 -0.0003 1.0000 0.0134 -8.500 -0.6990 0.07089 0.06747 -0.0116 1.0000 0.0130 -8.250 -0.7017 0.06152 0.05791 -0.0192 1.0000 0.0126 -8.000 -0.6997 0.05260 0.04867 -0.0239 1.0000 0.0123 -7.750 -0.6890 0.04704 0.04280 -0.0262 1.0000 0.0125 -7.500 -0.6717 0.04468 0.04023 -0.0273 1.0000 0.0135 -7.250 -0.6541 0.04033 0.03554 -0.0287 1.0000 0.0144 -7.000 -0.6350 0.03458 0.02924 -0.0299 1.0000 0.0143 -6.750 -0.6126 0.02921 0.02321 -0.0307 1.0000 0.0141 -6.500 -0.5876 0.02523 0.01862 -0.0311 1.0000 0.0140 -6.250 -0.5613 0.02240 0.01532 -0.0313 1.0000 0.0141 -6.000 -0.5347 0.02024 0.01281 -0.0313 1.0000 0.0144 -5.750 -0.5080 0.01851 0.01082 -0.0313 1.0000 0.0149 -5.500 -0.4811 0.01709 0.00920 -0.0312 1.0000 0.0156 -5.250 -0.4541 0.01589 0.00782 -0.0312 1.0000 0.0167 -5.000 -0.4267 0.01493 0.00672 -0.0313 1.0000 0.0178 -4.750 -0.3992 0.01400 0.00576 -0.0316 1.0000 0.0213 -4.500 -0.3715 0.01345 0.00515 -0.0317 1.0000 0.0268 -4.250 -0.3433 0.01269 0.00436 -0.0320 1.0000 0.0366 -4.000 -0.3153 0.01221 0.00384 -0.0322 1.0000 0.0492 -3.750 -0.2876 0.01181 0.00347 -0.0324 1.0000 0.0656 -3.500 -0.2597 0.01133 0.00309 -0.0328 1.0000 0.0913 -3.250 -0.2315 0.01078 0.00273 -0.0332 1.0000 0.1366 -3.000 -0.2030 0.00999 0.00240 -0.0341 1.0000 0.2532 -2.750 -0.1743 0.00885 0.00217 -0.0352 1.0000 0.4855 -2.500 -0.1477 0.00848 0.00218 -0.0349 1.0000 0.5974 -2.250 -0.1219 0.00831 0.00220 -0.0344 1.0000 0.6568 -2.000 -0.0967 0.00821 0.00224 -0.0336 1.0000 0.7049 -1.750 -0.0718 0.00815 0.00226 -0.0327 1.0000 0.7403 -1.500 -0.0475 0.00812 0.00231 -0.0317 1.0000 0.7699 -1.250 -0.0236 0.00810 0.00236 -0.0307 1.0000 0.7963 -1.000 0.0002 0.00808 0.00241 -0.0296 1.0000 0.8148 -0.750 0.0249 0.00807 0.00243 -0.0289 1.0000 0.8283 -0.500 0.0498 0.00806 0.00247 -0.0283 1.0000 0.8404 -0.250 0.0745 0.00807 0.00253 -0.0276 1.0000 0.8522 0.000 0.0993 0.00808 0.00261 -0.0270 1.0000 0.8640 0.250 0.1237 0.00810 0.00271 -0.0264 1.0000 0.8763 0.500 0.1708 0.00800 0.00269 -0.0304 0.9827 0.8849 0.750 0.2144 0.00788 0.00263 -0.0334 0.9491 0.8942 1.000 0.2503 0.00779 0.00256 -0.0343 0.8970 0.9051 1.250 0.2753 0.00782 0.00250 -0.0327 0.8209 0.9186 1.500 0.2944 0.00810 0.00240 -0.0298 0.7020 0.9349 1.750 0.3135 0.00892 0.00238 -0.0276 0.4975 0.9564 2.000 0.3390 0.01008 0.00261 -0.0278 0.2686 1.0000 2.250 0.3664 0.01099 0.00295 -0.0284 0.1380 1.0000 2.500 0.3949 0.01156 0.00334 -0.0289 0.0904 1.0000 2.750 0.4234 0.01206 0.00372 -0.0293 0.0630 1.0000 3.000 0.4518 0.01252 0.00409 -0.0297 0.0433 1.0000 3.250 0.4801 0.01303 0.00460 -0.0299 0.0313 1.0000 3.500 0.5078 0.01381 0.00538 -0.0300 0.0229 1.0000 3.750 0.5355 0.01457 0.00630 -0.0300 0.0203 1.0000 4.000 0.5626 0.01549 0.00732 -0.0299 0.0187 1.0000 4.250 0.5895 0.01654 0.00847 -0.0297 0.0178 1.0000 4.500 0.6163 0.01777 0.00982 -0.0295 0.0171 1.0000 4.750 0.6429 0.01897 0.01114 -0.0293 0.0159 1.0000 5.000 0.6689 0.02069 0.01311 -0.0291 0.0145 1.0000 5.250 0.6952 0.02250 0.01526 -0.0288 0.0139 1.0000 5.500 0.7205 0.02497 0.01817 -0.0282 0.0134 1.0000 5.750 0.7441 0.02823 0.02198 -0.0275 0.0131 1.0000 6.000 0.7651 0.03249 0.02686 -0.0267 0.0130 1.0000 6.250 0.7828 0.03775 0.03271 -0.0258 0.0130 1.0000 6.500 0.7976 0.04359 0.03907 -0.0252 0.0132 1.0000 6.750 0.8096 0.04962 0.04551 -0.0251 0.0134 1.0000 7.000 0.8189 0.05571 0.05193 -0.0255 0.0137 1.0000 7.250 0.8253 0.06181 0.05829 -0.0265 0.0139 1.0000 7.500 0.8285 0.06794 0.06463 -0.0282 0.0142 1.0000 7.750 0.8285 0.07408 0.07090 -0.0306 0.0144 1.0000 8.000 0.8258 0.08033 0.07726 -0.0339 0.0146 1.0000 8.250 0.7144 0.07578 0.07291 -0.0285 0.0145 1.0000