XFOIL Version 6.94 Calculated polar for: USA 45 AIRFOIL 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 ------- -------- --------- --------- -------- ------- ------- 1.000 0.4943 0.01043 0.00434 -0.0500 0.3733 0.9683 1.500 0.5831 0.01073 0.00454 -0.0567 0.3657 0.9788 2.000 0.6665 0.01110 0.00474 -0.0625 0.3575 0.9899 2.500 0.7524 0.01132 0.00484 -0.0691 0.3494 0.9974 3.000 0.8100 0.01142 0.00493 -0.0700 0.3437 1.0000 3.500 0.8479 0.01157 0.00495 -0.0668 0.3324 1.0000 4.000 0.8857 0.01150 0.00476 -0.0636 0.3124 1.0000 4.500 0.9232 0.01166 0.00478 -0.0603 0.2907 1.0000 5.000 0.9611 0.01192 0.00494 -0.0571 0.2693 1.0000 5.500 0.9985 0.01229 0.00521 -0.0537 0.2516 1.0000 6.000 1.0279 0.01313 0.00572 -0.0492 0.2022 1.0000 6.500 1.0248 0.01571 0.00756 -0.0397 0.0878 1.0000 7.500 1.0523 0.01829 0.00995 -0.0255 0.0041 1.0000 8.000 1.0759 0.01913 0.01088 -0.0204 0.0042 1.0000 8.500 1.0992 0.02019 0.01205 -0.0158 0.0043 1.0000 9.000 1.1205 0.02154 0.01354 -0.0115 0.0044 1.0000 9.500 1.1385 0.02330 0.01546 -0.0075 0.0046 1.0000 10.000 1.1554 0.02542 0.01773 -0.0043 0.0049 1.0000 10.500 1.1675 0.02827 0.02078 -0.0017 0.0053 1.0000 11.000 1.1711 0.03239 0.02512 0.0000 0.0056 1.0000 11.500 1.1651 0.03818 0.03115 0.0004 0.0058 1.0000 12.000 1.1471 0.04611 0.03934 -0.0007 0.0060 1.0000 12.500 1.1207 0.05562 0.04913 -0.0027 0.0060 1.0000 13.000 1.1144 0.06273 0.05642 -0.0040 0.0065 1.0000 13.500 1.0923 0.07206 0.06596 -0.0062 0.0067 1.0000 14.000 1.0703 0.08167 0.07577 -0.0085 0.0069 1.0000 14.500 1.0495 0.09150 0.08578 -0.0112 0.0071 1.0000 15.000 1.0307 0.10123 0.09567 -0.0139 0.0074 1.0000 15.500 1.0156 0.11051 0.10508 -0.0166 0.0076 1.0000 16.000 1.0177 0.11729 0.11197 -0.0184 0.0082 1.0000 16.500 1.0189 0.12381 0.11856 -0.0200 0.0095 1.0000 17.000 1.0472 0.12493 0.11964 -0.0188 0.0120 1.0000