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 = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.000 -0.4322 0.10731 0.10576 -0.0192 0.9999 0.0092 -12.750 -0.4345 0.10374 0.10220 -0.0191 0.9999 0.0092 -12.500 -0.4455 0.09722 0.09570 -0.0201 0.9998 0.0094 -12.250 -0.4464 0.09215 0.09064 -0.0219 0.9995 0.0097 -12.000 -0.4439 0.08779 0.08628 -0.0236 0.9990 0.0098 -11.750 -0.4413 0.08332 0.08182 -0.0256 0.9984 0.0099 -11.500 -0.4374 0.07928 0.07777 -0.0274 0.9977 0.0102 -11.250 -0.4350 0.07463 0.07313 -0.0297 0.9970 0.0103 -11.000 -0.4332 0.06980 0.06831 -0.0323 0.9962 0.0105 -10.750 -0.4319 0.06466 0.06317 -0.0353 0.9953 0.0107 -10.500 -0.4312 0.05944 0.05796 -0.0388 0.9941 0.0110 -10.250 -0.4343 0.05406 0.05259 -0.0424 0.9905 0.0111 -9.750 -0.4371 0.03036 0.02777 -0.0934 0.9875 0.0094 -9.500 -0.3963 0.02275 0.01943 -0.1020 0.9874 0.0094 -9.250 -0.3586 0.01861 0.01476 -0.1067 0.9874 0.0095 -9.000 -0.3218 0.01525 0.01096 -0.1103 0.9876 0.0097 -8.750 -0.2887 0.01385 0.00940 -0.1123 0.9875 0.0105 -8.500 -0.2557 0.01275 0.00816 -0.1140 0.9874 0.0110 -8.250 -0.2232 0.01199 0.00729 -0.1154 0.9871 0.0118 -8.000 -0.1907 0.01134 0.00656 -0.1168 0.9869 0.0127 -7.750 -0.1581 0.01075 0.00589 -0.1182 0.9867 0.0134 -7.500 -0.1254 0.01026 0.00532 -0.1195 0.9865 0.0139 -7.250 -0.0929 0.00985 0.00484 -0.1209 0.9863 0.0145 -7.000 -0.0576 0.00903 0.00390 -0.1229 0.9865 0.0179 -6.750 -0.0252 0.00879 0.00363 -0.1241 0.9862 0.0208 -6.500 0.0078 0.00850 0.00328 -0.1255 0.9859 0.0261 -6.250 0.0407 0.00826 0.00306 -0.1268 0.9857 0.0383 -6.000 0.0743 0.00789 0.00291 -0.1285 0.9856 0.0945 -5.750 0.1075 0.00757 0.00282 -0.1301 0.9856 0.1594 -5.500 0.1416 0.00717 0.00276 -0.1320 0.9858 0.2675 -5.250 0.1747 0.00693 0.00274 -0.1335 0.9857 0.3417 -5.000 0.2073 0.00676 0.00274 -0.1349 0.9856 0.4059 -4.750 0.2397 0.00665 0.00275 -0.1361 0.9854 0.4494 -4.500 0.2723 0.00652 0.00280 -0.1374 0.9852 0.5115 -4.250 0.3047 0.00645 0.00283 -0.1385 0.9848 0.5494 -4.000 0.3380 0.00638 0.00284 -0.1398 0.9844 0.5774 -3.750 0.3750 0.00624 0.00279 -0.1419 0.9836 0.6058 -3.500 0.4052 0.00605 0.00267 -0.1423 0.9789 0.6299 -3.250 0.4495 0.00563 0.00235 -0.1457 0.9756 0.6576 -3.000 0.4972 0.00505 0.00186 -0.1496 0.9702 0.6849 -2.750 0.5254 0.00478 0.00162 -0.1491 0.9575 0.7015 -2.500 0.5502 0.00466 0.00153 -0.1479 0.9406 0.7151 -2.250 0.5761 0.00461 0.00148 -0.1470 0.9197 0.7274 -2.000 0.6026 0.00462 0.00148 -0.1464 0.8950 0.7377 -1.750 0.6179 0.00566 0.00159 -0.1431 0.6756 0.7470 -1.500 0.6338 0.00740 0.00208 -0.1410 0.3735 0.7566 -1.250 0.6550 0.00861 0.00247 -0.1400 0.1596 0.7659 -1.000 0.6794 0.00933 0.00278 -0.1394 0.0545 0.7748 -0.750 0.7059 0.00970 0.00304 -0.1391 0.0292 0.7841 -0.500 0.7324 0.01004 0.00342 -0.1386 0.0216 0.7919 -0.250 0.7592 0.01031 0.00372 -0.1383 0.0183 0.7997 0.000 0.7840 0.01108 0.00461 -0.1374 0.0144 0.8079 0.250 0.8105 0.01139 0.00497 -0.1369 0.0139 0.8157 0.500 0.8365 0.01182 0.00546 -0.1364 0.0133 0.8235 0.750 0.8621 0.01231 0.00602 -0.1357 0.0125 0.8303 1.000 0.8876 0.01285 0.00662 -0.1351 0.0117 0.8380 1.250 0.9127 0.01343 0.00728 -0.1344 0.0109 0.8445 1.500 0.9375 0.01422 0.00814 -0.1335 0.0103 0.8518 1.750 0.9599 0.01588 0.00993 -0.1322 0.0095 0.8585 2.000 0.9828 0.01873 0.01306 -0.1307 0.0092 0.8660 2.250 1.0074 0.01980 0.01432 -0.1297 0.0090 0.8738 2.500 1.0322 0.02077 0.01548 -0.1288 0.0087 0.8817 2.750 1.0557 0.02237 0.01733 -0.1276 0.0084 0.8893 3.000 1.0768 0.02533 0.02069 -0.1255 0.0079 0.8966 7.750 1.0119 0.14049 0.13932 -0.1227 0.0080 1.0001 8.000 1.0107 0.14618 0.14500 -0.1255 0.0074 1.0001