XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC-SC2 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -14.000 -0.9596 0.06317 0.05971 -0.0483 1.0000 0.0161 -13.750 -0.9870 0.05446 0.05074 -0.0552 1.0000 0.0161 -13.500 -1.0075 0.04868 0.04474 -0.0583 1.0000 0.0161 -13.250 -1.0232 0.04448 0.04035 -0.0590 1.0000 0.0162 -13.000 -1.0343 0.04133 0.03702 -0.0583 1.0000 0.0163 -12.750 -1.0412 0.03891 0.03445 -0.0566 1.0000 0.0163 -12.500 -1.0459 0.03695 0.03235 -0.0541 1.0000 0.0164 -12.250 -1.0497 0.03535 0.03063 -0.0508 1.0000 0.0165 -12.000 -1.0538 0.03399 0.02914 -0.0468 1.0000 0.0165 -11.750 -1.0539 0.03266 0.02768 -0.0432 1.0000 0.0166 -11.500 -1.0523 0.03145 0.02634 -0.0396 1.0000 0.0167 -11.250 -1.0489 0.03031 0.02508 -0.0361 1.0000 0.0168 -11.000 -1.0437 0.02921 0.02386 -0.0328 1.0000 0.0169 -10.750 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0.4652 0.01067 0.00360 -0.0135 0.0722 0.9759 4.000 0.4972 0.01094 0.00380 -0.0145 0.0620 0.9784 4.250 0.5281 0.01121 0.00401 -0.0152 0.0527 0.9819 4.500 0.5592 0.01150 0.00424 -0.0160 0.0419 0.9849 4.750 0.5917 0.01184 0.00448 -0.0172 0.0329 0.9866 5.000 0.6231 0.01215 0.00476 -0.0181 0.0290 0.9887 5.250 0.6535 0.01247 0.00507 -0.0188 0.0268 0.9910 5.500 0.6831 0.01278 0.00538 -0.0193 0.0253 0.9933 5.750 0.7132 0.01312 0.00572 -0.0199 0.0240 0.9950 6.000 0.7433 0.01354 0.00613 -0.0206 0.0230 0.9966 6.250 0.7735 0.01388 0.00650 -0.0213 0.0224 0.9984 6.500 0.8035 0.01426 0.00690 -0.0219 0.0218 1.0000 6.750 0.8226 0.01460 0.00727 -0.0202 0.0214 1.0000 7.000 0.8415 0.01496 0.00765 -0.0185 0.0210 1.0000 7.250 0.8604 0.01534 0.00805 -0.0168 0.0206 1.0000 7.500 0.8791 0.01574 0.00848 -0.0150 0.0203 1.0000 7.750 0.8977 0.01617 0.00893 -0.0133 0.0200 1.0000 8.000 0.9159 0.01666 0.00944 -0.0115 0.0197 1.0000 8.250 0.9336 0.01722 0.01003 -0.0097 0.0194 1.0000 8.500 0.9504 0.01790 0.01074 -0.0077 0.0192 1.0000 8.750 0.9688 0.01843 0.01132 -0.0060 0.0190 1.0000 9.000 0.9874 0.01900 0.01194 -0.0044 0.0189 1.0000 9.250 1.0061 0.01958 0.01259 -0.0028 0.0187 1.0000 9.500 1.0250 0.02019 0.01326 -0.0013 0.0186 1.0000 9.750 1.0437 0.02083 0.01398 0.0001 0.0184 1.0000 10.000 1.0623 0.02151 0.01473 0.0016 0.0181 1.0000 10.250 1.0804 0.02225 0.01555 0.0030 0.0180 1.0000 10.500 1.0979 0.02304 0.01643 0.0046 0.0178 1.0000 10.750 1.1148 0.02388 0.01736 0.0061 0.0177 1.0000 11.000 1.1311 0.02476 0.01834 0.0077 0.0176 1.0000 11.250 1.1465 0.02570 0.01938 0.0094 0.0175 1.0000 11.500 1.1609 0.02668 0.02047 0.0112 0.0174 1.0000 11.750 1.1736 0.02769 0.02159 0.0132 0.0173 1.0000 12.000 1.1840 0.02875 0.02277 0.0155 0.0172 1.0000 12.250 1.1930 0.02989 0.02402 0.0180 0.0171 1.0000 12.500 1.2008 0.03110 0.02537 0.0203 0.0170 1.0000 12.750 1.2073 0.03243 0.02682 0.0227 0.0170 1.0000 13.000 1.2122 0.03388 0.02841 0.0250 0.0169 1.0000 13.250 1.2156 0.03547 0.03014 0.0273 0.0169 1.0000 13.500 1.2172 0.03724 0.03206 0.0294 0.0168 1.0000 13.750 1.2169 0.03924 0.03420 0.0313 0.0167 1.0000 14.000 1.2146 0.04151 0.03664 0.0328 0.0167 1.0000 14.250 1.2101 0.04414 0.03942 0.0340 0.0167 1.0000 14.500 1.2031 0.04724 0.04270 0.0345 0.0166 1.0000 14.750 1.1934 0.05099 0.04662 0.0342 0.0166 1.0000 15.000 1.1803 0.05561 0.05143 0.0328 0.0166 1.0000 15.250 1.1630 0.06155 0.05758 0.0298 0.0166 1.0000 15.500 1.1387 0.06962 0.06588 0.0245 0.0166 1.0000 15.750 1.0992 0.08224 0.07880 0.0155 0.0167 1.0000