XFOIL Version 6.96 Calculated polar for: NASA SC(2)-1006 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.750 -0.5375 0.08798 0.08580 -0.0190 0.9999 0.0166 -10.500 -0.5491 0.08573 0.08358 -0.0169 0.9999 0.0168 -10.250 -0.5595 0.08316 0.08106 -0.0156 0.9999 0.0170 -10.000 -0.5624 0.07955 0.07748 -0.0168 0.9999 0.0173 -9.750 -0.5621 0.07525 0.07320 -0.0199 0.9999 0.0177 -9.500 -0.5563 0.06965 0.06762 -0.0262 0.9999 0.0181 -9.250 -0.5405 0.06129 0.05920 -0.0381 0.9999 0.0187 -9.000 -0.5136 0.05169 0.04937 -0.0502 0.9999 0.0205 -8.750 -0.4786 0.04633 0.04363 -0.0568 0.9999 0.0220 -7.750 -0.2924 0.01759 0.01235 -0.0897 0.9999 0.0179 -7.500 -0.2535 0.01576 0.01023 -0.0923 0.9999 0.0186 -7.250 -0.2175 0.01455 0.00885 -0.0943 0.9999 0.0196 -7.000 -0.1856 0.01395 0.00815 -0.0954 0.9999 0.0211 -6.750 -0.1536 0.01343 0.00752 -0.0965 0.9999 0.0220 -6.500 -0.1046 0.01180 0.00573 -0.1015 0.9999 0.0260 -6.250 -0.0712 0.01143 0.00531 -0.1030 0.9999 0.0306 -6.000 -0.0362 0.01102 0.00487 -0.1048 0.9999 0.0413 -5.750 0.0041 0.01038 0.00450 -0.1080 0.9999 0.1174 -5.500 0.0474 0.00966 0.00436 -0.1122 0.9999 0.2819 -5.250 0.0833 0.00936 0.00437 -0.1144 0.9999 0.3871 -5.000 0.1167 0.00919 0.00442 -0.1159 0.9999 0.4613 -4.750 0.1482 0.00911 0.00450 -0.1170 0.9999 0.5218 -4.500 0.1791 0.00907 0.00462 -0.1178 0.9999 0.5789 -4.250 0.2082 0.00911 0.00476 -0.1182 0.9999 0.6168 -4.000 0.2365 0.00918 0.00491 -0.1184 0.9999 0.6454 -3.750 0.2642 0.00929 0.00507 -0.1185 0.9999 0.6720 -3.500 0.2916 0.00941 0.00526 -0.1185 0.9999 0.6954 -3.250 0.3192 0.00954 0.00546 -0.1186 0.9997 0.7190 -3.000 0.3545 0.00956 0.00556 -0.1202 0.9974 0.7417 -2.750 0.3992 0.00948 0.00555 -0.1238 0.9946 0.7580 -2.500 0.4440 0.00921 0.00535 -0.1273 0.9877 0.7717 -2.250 0.5023 0.00856 0.00477 -0.1334 0.9774 0.7838 -2.000 0.5752 0.00713 0.00346 -0.1419 0.9605 0.7935 -1.750 0.6138 0.00654 0.00293 -0.1433 0.9430 0.8032 -1.500 0.6418 0.00633 0.00277 -0.1426 0.9191 0.8132 -1.250 0.6645 0.00652 0.00243 -0.1401 0.7541 0.8217 -1.000 0.6696 0.00910 0.00313 -0.1356 0.3517 0.8299 -0.750 0.6866 0.01097 0.00375 -0.1340 0.0680 0.8394 -0.500 0.7103 0.01169 0.00435 -0.1329 0.0331 0.8481 -0.250 0.7361 0.01204 0.00477 -0.1322 0.0282 0.8566 0.000 0.7604 0.01280 0.00561 -0.1312 0.0237 0.8654 0.250 0.7831 0.01374 0.00667 -0.1298 0.0219 0.8731 0.500 0.8083 0.01418 0.00717 -0.1291 0.0201 0.8821 0.750 0.8322 0.01497 0.00805 -0.1279 0.0190 0.8917 1.000 0.8556 0.01585 0.00903 -0.1267 0.0179 0.9016 1.250 0.8789 0.01686 0.01015 -0.1254 0.0170 0.9128 1.500 0.9021 0.01821 0.01167 -0.1239 0.0166 0.9253 1.750 0.9256 0.02091 0.01472 -0.1218 0.0177 0.9410 2.000 0.9480 0.02254 0.01660 -0.1201 0.0172 1.0001 4.750 1.0361 0.05622 0.05389 -0.0854 0.0154 1.0001 5.000 1.0280 0.06136 0.05917 -0.0826 0.0151 1.0001 5.250 1.0104 0.06730 0.06527 -0.0793 0.0148 1.0001 5.500 0.9875 0.07234 0.07043 -0.0754 0.0147 1.0001 5.750 0.9691 0.07568 0.07388 -0.0720 0.0148 1.0001 6.000 0.9463 0.08006 0.07837 -0.0699 0.0149 1.0001 6.250 0.9237 0.08547 0.08388 -0.0700 0.0150 1.0001