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 = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.250 -0.5059 0.09792 0.09430 -0.0227 0.9999 0.0399 -11.000 -0.5100 0.09497 0.09140 -0.0229 0.9999 0.0414 -10.750 -0.5182 0.09221 0.08871 -0.0229 0.9999 0.0428 -10.500 -0.4466 0.07688 0.07368 -0.0238 0.9999 0.0479 -10.250 -0.4511 0.07399 0.07082 -0.0226 0.9999 0.0490 -10.000 -0.5358 0.08060 0.07729 -0.0343 0.9999 0.0446 -9.750 -0.5285 0.07410 0.07066 -0.0435 0.9999 0.0448 -9.500 -0.5333 0.06979 0.06651 -0.0388 0.9999 0.0467 -9.250 -0.5275 0.06900 0.06575 -0.0346 0.9999 0.0492 -9.000 -0.5139 0.06364 0.06033 -0.0407 0.9999 0.0515 -8.750 -0.4722 0.05143 0.04731 -0.0629 0.9999 0.0585 -8.500 -0.4590 0.04781 0.04389 -0.0621 0.9999 0.0604 -8.250 -0.4369 0.04475 0.04073 -0.0640 0.9999 0.0639 -8.000 -0.3958 0.03873 0.03417 -0.0725 0.9999 0.0738 -7.750 -0.3298 0.02801 0.02218 -0.0822 0.9999 0.0422 -7.500 -0.2831 0.02316 0.01661 -0.0872 0.9999 0.0386 -7.250 -0.2441 0.02066 0.01363 -0.0898 0.9999 0.0394 -7.000 -0.2045 0.01844 0.01101 -0.0921 0.9999 0.0392 -6.750 -0.1681 0.01694 0.00928 -0.0939 0.9999 0.0411 -6.500 -0.1347 0.01611 0.00830 -0.0950 0.9999 0.0448 -6.250 -0.0941 0.01472 0.00694 -0.0981 0.9999 0.0548 -6.000 -0.0535 0.01373 0.00597 -0.1013 0.9999 0.0743 -5.750 0.0066 0.01170 0.00523 -0.1099 0.9999 0.3901 -5.500 0.0397 0.01142 0.00527 -0.1113 0.9999 0.5019 -5.250 0.0689 0.01133 0.00540 -0.1116 0.9999 0.5817 -5.000 0.0958 0.01136 0.00556 -0.1112 0.9999 0.6363 -4.750 0.1218 0.01145 0.00568 -0.1107 0.9999 0.6742 -4.500 0.1473 0.01155 0.00581 -0.1101 0.9999 0.7039 -4.250 0.1725 0.01168 0.00596 -0.1095 0.9999 0.7303 -4.000 0.1967 0.01182 0.00614 -0.1085 0.9999 0.7567 -3.750 0.2197 0.01198 0.00635 -0.1073 0.9999 0.7813 -3.500 0.2428 0.01212 0.00652 -0.1061 0.9999 0.8020 -3.250 0.2665 0.01224 0.00668 -0.1052 0.9999 0.8209 -3.000 0.2912 0.01236 0.00683 -0.1046 0.9999 0.8391 -2.750 0.3144 0.01244 0.00698 -0.1036 0.9999 0.8554 -2.500 0.3377 0.01251 0.00712 -0.1026 0.9999 0.8724 -2.250 0.3615 0.01258 0.00726 -0.1018 0.9999 0.8909 -2.000 0.3841 0.01259 0.00736 -0.1008 0.9999 0.9104 -1.750 0.4051 0.01257 0.00742 -0.0995 0.9999 0.9341 -1.500 0.4152 0.01259 0.00755 -0.0962 0.9999 1.0001 -1.250 0.4473 0.01288 0.00794 -0.0977 0.9999 1.0001 -1.000 0.4783 0.01320 0.00834 -0.0990 0.9999 1.0001 -0.750 0.5754 0.01204 0.00741 -0.1123 0.9672 1.0001 -0.500 0.6494 0.00976 0.00542 -0.1188 0.9144 1.0001 -0.250 0.6714 0.01305 0.00484 -0.1155 0.0969 1.0001 0.000 0.6974 0.01444 0.00600 -0.1152 0.0571 1.0001 0.250 0.7238 0.01565 0.00717 -0.1149 0.0458 1.0001 0.500 0.7515 0.01670 0.00830 -0.1147 0.0416 1.0001 0.750 0.7794 0.01803 0.00967 -0.1144 0.0387 1.0001 1.000 0.8084 0.01963 0.01133 -0.1142 0.0372 1.0001 1.250 0.8375 0.02149 0.01322 -0.1143 0.0348 1.0001 1.500 0.8671 0.02521 0.01719 -0.1142 0.0333 1.0001 1.750 0.8962 0.02691 0.01917 -0.1137 0.0339 1.0001 2.000 0.9262 0.03137 0.02441 -0.1116 0.0436 1.0001 3.250 1.0301 0.05026 0.04561 -0.1014 0.0507 1.0001 3.500 1.0453 0.05356 0.04900 -0.1003 0.0480 1.0001 3.750 1.0585 0.05908 0.05440 -0.0999 0.0462 1.0001 4.000 1.0567 0.06641 0.06214 -0.0975 0.0451 1.0001 4.250 1.0666 0.06771 0.06422 -0.0941 0.0406 1.0001 4.500 1.0710 0.07216 0.06889 -0.0926 0.0386 1.0001 4.750 1.0730 0.07650 0.07340 -0.0913 0.0372 1.0001 5.000 1.0730 0.08074 0.07776 -0.0900 0.0362 1.0001 5.250 1.0715 0.08488 0.08200 -0.0888 0.0354 1.0001 5.500 1.0693 0.08898 0.08615 -0.0876 0.0347 1.0001 5.750 1.0658 0.09348 0.09067 -0.0865 0.0341 1.0001 6.000 1.0540 0.10094 0.09811 -0.0860 0.0334 1.0001 6.250 0.8580 0.10306 0.10065 -0.0794 0.0418 1.0001