XFOIL Version 6.96 Calculated polar for: NASA SC(2)-1010 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 ------ -------- --------- --------- -------- -------- -------- -17.250 -0.8842 0.07398 0.07159 -0.0572 0.9999 0.0148 -17.000 -0.9250 0.06216 0.05957 -0.0654 0.9999 0.0147 -16.750 -0.9487 0.05422 0.05146 -0.0711 0.9999 0.0147 -16.500 -0.9657 0.04806 0.04514 -0.0755 0.9999 0.0147 -16.250 -0.9782 0.04336 0.04031 -0.0783 0.9999 0.0148 -16.000 -0.9899 0.03971 0.03655 -0.0794 0.9999 0.0149 -15.750 -1.0068 0.03695 0.03369 -0.0781 0.9999 0.0149 -15.500 -1.0320 0.03498 0.03166 -0.0740 0.9999 0.0149 -15.250 -1.0384 0.03275 0.02933 -0.0735 0.9999 0.0150 -15.000 -1.0223 0.03060 0.02702 -0.0767 0.9992 0.0151 -14.750 -0.9929 0.02860 0.02484 -0.0816 0.9979 0.0153 -14.500 -0.9681 0.02467 0.02059 -0.0873 0.9965 0.0156 -14.250 -0.9384 0.02266 0.01843 -0.0910 0.9956 0.0160 -14.000 -0.9069 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0.8458 0.01082 0.00399 -0.1461 0.0753 0.7154 0.750 0.8989 0.01147 0.00450 -0.1454 0.0491 0.7289 1.000 0.9259 0.01172 0.00476 -0.1451 0.0446 0.7353 1.250 0.9525 0.01202 0.00503 -0.1447 0.0405 0.7410 1.500 0.9790 0.01230 0.00533 -0.1443 0.0377 0.7467 1.750 1.0057 0.01255 0.00559 -0.1439 0.0356 0.7526 2.000 1.0316 0.01291 0.00594 -0.1434 0.0329 0.7579 2.250 1.0575 0.01323 0.00631 -0.1429 0.0311 0.7634 2.500 1.0839 0.01350 0.00658 -0.1424 0.0295 0.7693 2.750 1.1097 0.01381 0.00690 -0.1419 0.0278 0.7745 3.000 1.1342 0.01430 0.00742 -0.1411 0.0259 0.7798 3.250 1.1602 0.01458 0.00772 -0.1406 0.0249 0.7850 3.500 1.1857 0.01490 0.00807 -0.1400 0.0238 0.7895 3.750 1.2106 0.01527 0.00845 -0.1393 0.0227 0.7940 4.000 1.2328 0.01603 0.00925 -0.1381 0.0214 0.7987 4.250 1.2580 0.01636 0.00962 -0.1374 0.0209 0.8034 4.500 1.2822 0.01676 0.01008 -0.1366 0.0203 0.8081 4.750 1.3062 0.01720 0.01056 -0.1357 0.0197 0.8136 5.000 1.3300 0.01763 0.01102 -0.1348 0.0191 0.8188 5.250 1.3534 0.01807 0.01150 -0.1338 0.0185 0.8234 5.500 1.3742 0.01886 0.01232 -0.1324 0.0178 0.8282 5.750 1.3943 0.01979 0.01332 -0.1309 0.0173 0.8326 6.000 1.4168 0.02028 0.01389 -0.1298 0.0170 0.8365 6.250 1.4387 0.02084 0.01453 -0.1286 0.0167 0.8407 6.500 1.4601 0.02147 0.01523 -0.1273 0.0164 0.8453 6.750 1.4810 0.02211 0.01593 -0.1259 0.0160 0.8495 7.000 1.5012 0.02276 0.01666 -0.1245 0.0157 0.8541 7.250 1.5210 0.02344 0.01740 -0.1229 0.0154 0.8593 7.500 1.5403 0.02411 0.01814 -0.1214 0.0151 0.8640 7.750 1.5583 0.02487 0.01897 -0.1196 0.0149 0.8686 8.000 1.5750 0.02575 0.01991 -0.1176 0.0146 0.8737 8.250 1.5882 0.02699 0.02126 -0.1151 0.0143 0.8786 8.500 1.5939 0.02923 0.02371 -0.1116 0.0140 0.8838 8.750 1.6068 0.03016 0.02477 -0.1090 0.0139 0.8895 9.000 1.6187 0.03098 0.02571 -0.1063 0.0137 0.8957 9.250 1.6273 0.03181 0.02666 -0.1030 0.0136 0.9033 9.500 1.6346 0.03289 0.02789 -0.0997 0.0135 0.9132 9.750 1.6398 0.03407 0.02925 -0.0962 0.0133 0.9327 10.000 1.6434 0.03534 0.03068 -0.0926 0.0132 1.0001 10.250 1.6483 0.03688 0.03235 -0.0896 0.0131 1.0001 10.500 1.6515 0.03857 0.03418 -0.0865 0.0129 1.0001 10.750 1.6533 0.04042 0.03616 -0.0835 0.0128 1.0001 11.000 1.6542 0.04236 0.03824 -0.0807 0.0127 1.0001 11.250 1.6533 0.04452 0.04054 -0.0779 0.0125 1.0001 11.500 1.6498 0.04705 0.04321 -0.0753 0.0124 1.0001 11.750 1.6463 0.04960 0.04590 -0.0730 0.0123 1.0001 12.000 1.6400 0.05262 0.04906 -0.0709 0.0123 1.0001 12.250 1.6319 0.05598 0.05257 -0.0692 0.0122 1.0001 12.500 1.6218 0.05973 0.05648 -0.0679 0.0121 1.0001 12.750 1.6107 0.06389 0.06079 -0.0673 0.0120 1.0001 13.000 1.5956 0.06891 0.06597 -0.0674 0.0120 1.0001 13.250 1.5786 0.07466 0.07189 -0.0686 0.0120 1.0001 13.500 1.5553 0.08218 0.07961 -0.0713 0.0119 1.0001 13.750 1.5247 0.09214 0.08980 -0.0765 0.0120 1.0001 14.000 1.4815 0.10622 0.10417 -0.0855 0.0120 1.0001