XFOIL Version 6.94 Calculated polar for: SA7036 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 0.3471 0.00573 0.00175 -0.0708 0.7762 0.9899 0.500 0.4070 0.00584 0.00163 -0.0716 0.7400 1.0000 1.000 0.4570 0.00605 0.00160 -0.0702 0.7023 1.0000 1.500 0.5082 0.00632 0.00165 -0.0691 0.6643 1.0000 2.000 0.5600 0.00661 0.00174 -0.0681 0.6225 1.0000 2.500 0.6120 0.00695 0.00188 -0.0673 0.5806 1.0000 3.000 0.6639 0.00733 0.00209 -0.0665 0.5392 1.0000 3.500 0.7160 0.00772 0.00233 -0.0658 0.4972 1.0000 4.000 0.7676 0.00817 0.00265 -0.0650 0.4536 1.0000 4.500 0.8188 0.00868 0.00300 -0.0642 0.4087 1.0000 5.000 0.8661 0.00953 0.00342 -0.0630 0.3237 1.0000 5.500 0.9149 0.01030 0.00396 -0.0620 0.2637 1.0000 6.000 0.9604 0.01143 0.00462 -0.0607 0.1805 1.0000 6.500 1.0071 0.01243 0.00537 -0.0595 0.1274 1.0000 7.000 1.0507 0.01374 0.00638 -0.0579 0.0703 1.0000 7.500 1.0889 0.01558 0.00790 -0.0555 0.0130 1.0000 8.000 1.1326 0.01675 0.00917 -0.0536 0.0063 1.0000 8.500 1.1751 0.01795 0.01057 -0.0516 0.0057 1.0000 9.000 1.2140 0.01939 0.01225 -0.0491 0.0055 1.0000 9.500 1.2464 0.02120 0.01432 -0.0457 0.0054 1.0000 10.000 1.2671 0.02343 0.01688 -0.0406 0.0054 1.0000 10.500 1.2759 0.02605 0.01979 -0.0343 0.0055 1.0000 11.000 1.2786 0.02945 0.02351 -0.0285 0.0056 1.0000 11.500 1.2775 0.03377 0.02817 -0.0236 0.0058 1.0000 12.000 1.2741 0.03898 0.03375 -0.0201 0.0060 1.0000 12.500 1.2652 0.04545 0.04063 -0.0179 0.0062 1.0000 13.000 1.2488 0.05351 0.04912 -0.0177 0.0064 1.0000 13.500 1.2254 0.06337 0.05940 -0.0198 0.0066 1.0000 14.000 1.1964 0.07528 0.07170 -0.0248 0.0067 1.0000 14.500 1.1648 0.08925 0.08603 -0.0322 0.0068 1.0000 15.000 1.1314 0.10530 0.10240 -0.0418 0.0069 1.0000 15.500 1.0961 0.12370 0.12109 -0.0532 0.0070 1.0000 16.000 1.0618 0.14367 0.14130 -0.0655 0.0070 1.0000 16.500 1.0236 0.16727 0.16510 -0.0793 0.0072 1.0000