XFOIL Version 6.94 Calculated polar for: EPPLER 544 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2691 0.01614 0.01010 -0.0494 0.6909 0.7518 0.500 0.3244 0.01597 0.00982 -0.0500 0.6839 0.7539 1.000 0.3731 0.01578 0.00969 -0.0493 0.6762 0.7560 1.500 0.4264 0.01566 0.00956 -0.0495 0.6683 0.7589 2.000 0.4822 0.01557 0.00941 -0.0504 0.6606 0.7615 2.500 0.5321 0.01533 0.00923 -0.0502 0.6512 0.7644 3.000 0.5898 0.01515 0.00900 -0.0516 0.6428 0.7672 3.500 0.6420 0.01511 0.00900 -0.0519 0.6330 0.7705 4.000 0.6966 0.01484 0.00876 -0.0523 0.6227 0.7726 4.500 0.7470 0.01477 0.00879 -0.0518 0.6120 0.7747 5.000 0.8001 0.01461 0.00868 -0.0517 0.5999 0.7771 5.500 0.8479 0.01454 0.00875 -0.0507 0.5862 0.7797 6.000 0.8992 0.01445 0.00871 -0.0503 0.5717 0.7826 6.500 0.9465 0.01434 0.00867 -0.0491 0.5540 0.7864 7.000 0.9877 0.01425 0.00869 -0.0467 0.5325 0.7896 7.500 1.0224 0.01422 0.00878 -0.0432 0.5063 0.7925 8.000 1.0440 0.01427 0.00891 -0.0369 0.4717 0.7959 8.500 1.0514 0.01475 0.00931 -0.0283 0.4264 0.7999 9.000 1.0523 0.01587 0.01026 -0.0195 0.3732 0.8046 9.500 1.0478 0.01754 0.01172 -0.0110 0.3222 0.8099 10.000 1.0423 0.01955 0.01358 -0.0033 0.2763 0.8150 10.500 1.0379 0.02194 0.01587 0.0033 0.2340 0.8204 11.000 1.0366 0.02465 0.01848 0.0085 0.1953 0.8264 12.000 1.0390 0.03071 0.02442 0.0163 0.1306 0.8392 12.500 1.0411 0.03416 0.02783 0.0193 0.1043 0.8472 13.000 1.0432 0.03786 0.03154 0.0217 0.0829 0.8562 13.500 1.0450 0.04176 0.03549 0.0238 0.0661 0.8687 14.000 1.0477 0.04589 0.03974 0.0254 0.0535 0.8853 14.500 1.0532 0.05068 0.04472 0.0252 0.0437 0.9233 15.000 1.0650 0.05595 0.05014 0.0227 0.0358 1.0000 15.500 1.0672 0.06171 0.05595 0.0217 0.0299 1.0000 16.000 1.0694 0.06765 0.06203 0.0205 0.0251 1.0000 16.500 1.0688 0.07417 0.06868 0.0189 0.0214 1.0000 17.000 1.0665 0.08121 0.07586 0.0168 0.0182 1.0000 17.500 1.0620 0.08881 0.08360 0.0142 0.0159 1.0000 18.000 1.0592 0.09648 0.09144 0.0112 0.0141 1.0000 19.000 1.0514 0.11266 0.10798 0.0041 0.0115 1.0000 19.500 1.0472 0.12092 0.11643 0.0001 0.0106 1.0000 20.000 1.0387 0.13043 0.12624 -0.0051 0.0099 1.0000 20.500 1.0323 0.13951 0.13554 -0.0103 0.0094 1.0000 21.000 1.0294 0.14770 0.14381 -0.0152 0.0088 1.0000 21.500 1.0113 0.15948 0.15590 -0.0224 0.0086 1.0000 22.000 0.9942 0.17133 0.16802 -0.0298 0.0086 1.0000 22.500 0.9647 0.18674 0.18374 -0.0395 0.0086 1.0000