XFOIL Version 6.94 Calculated polar for: N-9 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4619 0.00656 0.00170 -0.0649 0.6647 1.0000 0.500 0.5155 0.00676 0.00179 -0.0642 0.6352 1.0000 1.000 0.5687 0.00699 0.00189 -0.0636 0.5978 1.0000 1.500 0.6210 0.00731 0.00201 -0.0628 0.5381 1.0000 2.000 0.6704 0.00798 0.00224 -0.0616 0.4411 1.0000 2.500 0.7207 0.00869 0.00265 -0.0608 0.3750 1.0000 3.000 0.7724 0.00928 0.00307 -0.0602 0.3332 1.0000 3.500 0.8241 0.00987 0.00351 -0.0597 0.2881 1.0000 4.000 0.8703 0.01115 0.00417 -0.0585 0.1549 1.0000 4.500 0.9180 0.01227 0.00512 -0.0575 0.1220 1.0000 5.000 0.9670 0.01318 0.00605 -0.0566 0.1106 1.0000 5.500 1.0156 0.01407 0.00700 -0.0557 0.0993 1.0000 6.000 1.0644 0.01489 0.00789 -0.0549 0.0891 1.0000 6.500 1.1143 0.01552 0.00860 -0.0543 0.0786 1.0000 7.000 1.1655 0.01597 0.00913 -0.0540 0.0654 1.0000 7.500 1.2085 0.01737 0.01025 -0.0525 0.0257 1.0000 8.000 1.2452 0.01941 0.01240 -0.0499 0.0173 1.0000 8.500 1.2796 0.02146 0.01465 -0.0471 0.0145 1.0000 9.000 1.3020 0.02440 0.01785 -0.0428 0.0130 1.0000 9.500 1.3261 0.02675 0.02044 -0.0388 0.0118 1.0000 10.000 1.3384 0.02951 0.02340 -0.0336 0.0109 1.0000 10.500 1.3406 0.03346 0.02758 -0.0286 0.0103 1.0000 11.000 1.3324 0.03931 0.03377 -0.0241 0.0099 1.0000 11.500 1.3331 0.04425 0.03909 -0.0219 0.0097 1.0000 12.000 1.3266 0.05063 0.04586 -0.0213 0.0094 1.0000 12.500 1.3128 0.05862 0.05424 -0.0225 0.0092 1.0000 13.000 1.2916 0.06842 0.06440 -0.0256 0.0090 1.0000 13.500 1.2647 0.08008 0.07642 -0.0306 0.0089 1.0000 14.000 1.2323 0.09417 0.09085 -0.0378 0.0089 1.0000 14.500 1.1954 0.11052 0.10753 -0.0468 0.0090 1.0000 15.000 1.1539 0.12907 0.12637 -0.0572 0.0093 1.0000 15.500 1.1118 0.14962 0.14717 -0.0690 0.0096 1.0000