XFOIL Version 6.96 Calculated polar for: NACA 16-006 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.000 -0.5185 0.08936 0.08473 -0.0124 1.0000 0.0402 -8.750 -0.5228 0.08503 0.08043 -0.0138 1.0000 0.0408 -8.500 -0.5284 0.08061 0.07605 -0.0155 1.0000 0.0412 -8.250 -0.5359 0.07602 0.07149 -0.0177 1.0000 0.0416 -8.000 -0.5463 0.07142 0.06692 -0.0204 1.0000 0.0416 -7.750 -0.5578 0.06730 0.06280 -0.0212 1.0000 0.0417 -7.500 -0.5666 0.06315 0.05863 -0.0214 1.0000 0.0419 -7.250 -0.5725 0.05886 0.05427 -0.0214 1.0000 0.0427 -7.000 -0.5763 0.05458 0.04988 -0.0211 1.0000 0.0436 -6.000 -0.6046 0.05050 0.04457 -0.0131 1.0000 0.0237 -5.750 -0.5946 0.04647 0.04027 -0.0111 1.0000 0.0210 -5.250 -0.5636 0.03910 0.03198 -0.0059 1.0000 0.0166 -5.000 -0.5497 0.03573 0.02821 -0.0036 1.0000 0.0162 -4.750 -0.5334 0.03262 0.02468 -0.0014 1.0000 0.0160 -4.500 -0.5148 0.02973 0.02133 0.0008 1.0000 0.0159 -4.250 -0.4941 0.02706 0.01818 0.0028 1.0000 0.0159 -4.000 -0.4711 0.02483 0.01548 0.0045 1.0000 0.0166 -3.750 -0.4484 0.02253 0.01287 0.0058 1.0000 0.0188 -3.500 -0.4245 0.02072 0.01085 0.0070 1.0000 0.0205 -3.250 -0.4011 0.01908 0.00904 0.0084 1.0000 0.0224 -3.000 -0.3792 0.01784 0.00760 0.0099 1.0000 0.0271 -2.750 -0.3588 0.01666 0.00632 0.0115 1.0000 0.0340 -2.500 -0.3377 0.01558 0.00519 0.0129 1.0000 0.0531 -2.250 -0.1480 0.01275 0.00501 -0.0172 1.0000 1.0000 -2.000 -0.1300 0.01259 0.00467 -0.0157 1.0000 1.0000 -1.750 -0.1125 0.01245 0.00434 -0.0141 1.0000 1.0000 -1.500 -0.0956 0.01233 0.00410 -0.0123 1.0000 1.0000 -1.250 -0.0791 0.01223 0.00391 -0.0104 1.0000 1.0000 -1.000 -0.0629 0.01216 0.00376 -0.0084 1.0000 1.0000 -0.750 -0.0469 0.01210 0.00365 -0.0064 1.0000 1.0000 -0.500 -0.0311 0.01206 0.00356 -0.0043 1.0000 1.0000 -0.250 -0.0155 0.01203 0.00351 -0.0022 1.0000 1.0000 0.000 0.0000 0.01202 0.00349 0.0000 1.0000 1.0000 0.250 0.0155 0.01203 0.00351 0.0022 1.0000 1.0000 0.500 0.0311 0.01206 0.00356 0.0043 1.0000 1.0000 0.750 0.0469 0.01210 0.00365 0.0064 1.0000 1.0000 1.000 0.0629 0.01216 0.00376 0.0084 1.0000 1.0000 1.250 0.0791 0.01223 0.00391 0.0104 1.0000 1.0000 1.500 0.0956 0.01233 0.00410 0.0123 1.0000 1.0000 1.750 0.1125 0.01245 0.00434 0.0141 1.0000 1.0000 2.000 0.1299 0.01259 0.00467 0.0157 1.0000 1.0000 2.250 0.1479 0.01275 0.00501 0.0172 1.0000 1.0000 2.500 0.3378 0.01558 0.00519 -0.0129 0.0530 1.0000 2.750 0.3589 0.01666 0.00632 -0.0115 0.0340 1.0000 3.000 0.3794 0.01784 0.00760 -0.0099 0.0271 1.0000 3.250 0.4012 0.01909 0.00900 -0.0084 0.0224 1.0000 3.500 0.4246 0.02072 0.01085 -0.0070 0.0205 1.0000 3.750 0.4486 0.02254 0.01288 -0.0058 0.0188 1.0000 4.000 0.4713 0.02484 0.01549 -0.0045 0.0166 1.0000 4.250 0.4942 0.02707 0.01820 -0.0028 0.0159 1.0000 4.500 0.5149 0.02973 0.02134 -0.0008 0.0159 1.0000 4.750 0.5335 0.03262 0.02468 0.0014 0.0160 1.0000 5.000 0.5498 0.03574 0.02822 0.0036 0.0162 1.0000 5.250 0.5636 0.03910 0.03199 0.0058 0.0166 1.0000 5.750 0.5946 0.04647 0.04028 0.0111 0.0210 1.0000 6.000 0.6046 0.05049 0.04457 0.0131 0.0237 1.0000 6.500 0.6246 0.06277 0.05690 0.0159 0.0460 1.0000 6.750 0.6323 0.06373 0.05854 0.0170 0.0413 1.0000 7.000 0.6378 0.06753 0.06248 0.0173 0.0398 1.0000 7.250 0.6421 0.07124 0.06627 0.0176 0.0382 1.0000 7.500 0.5664 0.06299 0.05846 0.0215 0.0422 1.0000 7.750 0.5573 0.06716 0.06266 0.0213 0.0416 1.0000 8.000 0.5459 0.07125 0.06675 0.0204 0.0416 1.0000 8.250 0.5354 0.07586 0.07133 0.0178 0.0416 1.0000 8.500 0.5280 0.08042 0.07586 0.0156 0.0413 1.0000 8.750 0.5223 0.08486 0.08027 0.0138 0.0408 1.0000 9.000 0.5180 0.08916 0.08453 0.0124 0.0403 1.0000