XFOIL Version 6.94 Calculated polar for: S4053 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 0.4012 0.00619 0.00184 -0.0840 0.7500 1.0000 0.500 0.4493 0.00630 0.00179 -0.0823 0.7316 1.0000 1.000 0.4982 0.00644 0.00180 -0.0807 0.7137 1.0000 1.500 0.5475 0.00659 0.00184 -0.0792 0.6954 1.0000 2.000 0.5970 0.00675 0.00192 -0.0778 0.6756 1.0000 2.500 0.6465 0.00690 0.00203 -0.0763 0.6540 1.0000 3.000 0.6956 0.00710 0.00215 -0.0748 0.6300 1.0000 3.500 0.7445 0.00730 0.00232 -0.0732 0.6002 1.0000 4.000 0.7927 0.00757 0.00251 -0.0715 0.5641 1.0000 4.500 0.8397 0.00794 0.00279 -0.0696 0.5154 1.0000 5.000 0.8828 0.00861 0.00316 -0.0671 0.4367 1.0000 5.500 0.9245 0.00952 0.00373 -0.0646 0.3505 1.0000 6.000 0.9650 0.01060 0.00445 -0.0621 0.2653 1.0000 7.000 1.0428 0.01313 0.00633 -0.0568 0.1190 1.0000 7.500 1.0802 0.01450 0.00747 -0.0539 0.0765 1.0000 8.000 1.1200 0.01559 0.00859 -0.0514 0.0568 1.0000 8.500 1.1572 0.01680 0.00987 -0.0485 0.0457 1.0000 9.000 1.1896 0.01828 0.01142 -0.0449 0.0358 1.0000 9.500 1.2170 0.01989 0.01310 -0.0406 0.0252 1.0000 10.000 1.2455 0.02116 0.01450 -0.0365 0.0187 1.0000 12.000 1.2763 0.03397 0.02828 -0.0158 0.0105 1.0000 12.500 1.2872 0.03746 0.03207 -0.0131 0.0099 1.0000 13.000 1.2868 0.04258 0.03754 -0.0107 0.0094 1.0000 13.500 1.2811 0.04858 0.04394 -0.0093 0.0090 1.0000 14.000 1.2693 0.05571 0.05143 -0.0093 0.0085 1.0000 14.500 1.2480 0.06486 0.06096 -0.0111 0.0083 1.0000 15.000 1.2196 0.07627 0.07275 -0.0153 0.0083 1.0000 15.500 1.1844 0.09055 0.08742 -0.0223 0.0083 1.0000 16.000 1.1419 0.10846 0.10571 -0.0327 0.0085 1.0000 16.500 1.0904 0.13116 0.12872 -0.0466 0.0089 1.0000