XFOIL Version 6.96 Calculated polar for: USA 50 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.4965 0.08623 0.08469 -0.0158 1.0000 0.0026 -8.000 -0.5014 0.08268 0.08116 -0.0166 1.0000 0.0025 -7.750 -0.5103 0.07968 0.07819 -0.0167 1.0000 0.0025 -7.250 -0.5200 0.07108 0.06962 -0.0212 1.0000 0.0023 -6.500 -0.4587 0.05055 0.04879 -0.0427 0.9939 0.0031 -6.250 -0.4379 0.04416 0.04221 -0.0470 0.9907 0.0028 -6.000 -0.4133 0.03757 0.03535 -0.0503 0.9882 0.0027 -5.750 -0.3937 0.03122 0.02866 -0.0509 0.9841 0.0025 -5.500 -0.3824 0.01714 0.01339 -0.0481 0.9792 0.0021 -5.250 -0.3620 0.01410 0.00983 -0.0465 0.9752 0.0021 -5.000 -0.3376 0.01230 0.00769 -0.0458 0.9715 0.0021 -4.750 -0.3089 0.01118 0.00636 -0.0460 0.9689 0.0022 -4.500 -0.2855 0.01048 0.00554 -0.0449 0.9628 0.0026 -4.250 -0.2589 0.00974 0.00467 -0.0446 0.9574 0.0029 -4.000 -0.2333 0.00912 0.00393 -0.0441 0.9511 0.0031 -3.750 -0.2043 0.00859 0.00330 -0.0443 0.9456 0.0035 -3.500 -0.1708 0.00823 0.00293 -0.0457 0.9411 0.0044 -3.250 -0.1360 0.00783 0.00245 -0.0473 0.9345 0.0053 -3.000 -0.0937 0.00741 0.00192 -0.0506 0.9282 0.0063 -2.750 -0.0525 0.00711 0.00155 -0.0537 0.9166 0.0092 -2.250 0.0149 0.00669 0.00110 -0.0565 0.8818 0.0487 -2.000 0.0425 0.00665 0.00100 -0.0564 0.8549 0.0609 -1.750 0.0662 0.00670 0.00090 -0.0554 0.8075 0.0725 -1.500 0.0864 0.00685 0.00082 -0.0536 0.7453 0.0853 -1.250 0.1084 0.00693 0.00076 -0.0524 0.7004 0.1038 -1.000 0.1304 0.00687 0.00070 -0.0512 0.6585 0.1629 -0.750 0.1512 0.00666 0.00065 -0.0499 0.6139 0.2780 -0.500 0.1731 0.00655 0.00064 -0.0488 0.5756 0.3650 -0.250 0.1955 0.00631 0.00063 -0.0479 0.5552 0.4723 0.000 0.2176 0.00603 0.00065 -0.0468 0.5421 0.5881 0.250 0.2393 0.00578 0.00067 -0.0455 0.5325 0.6885 0.500 0.2563 0.00535 0.00070 -0.0430 0.5235 0.8363 0.750 0.3171 0.00524 0.00084 -0.0506 0.5128 0.9462 1.000 0.3627 0.00536 0.00091 -0.0547 0.5012 0.9665 1.250 0.3956 0.00549 0.00099 -0.0559 0.4893 0.9797 1.500 0.4266 0.00561 0.00106 -0.0567 0.4743 0.9873 1.750 0.4640 0.00573 0.00113 -0.0590 0.4537 0.9921 2.000 0.4944 0.00591 0.00120 -0.0597 0.4197 0.9949 2.250 0.5236 0.00654 0.00134 -0.0605 0.3060 0.9977 2.500 0.5562 0.00701 0.00151 -0.0619 0.2331 1.0000 2.750 0.5714 0.00793 0.00185 -0.0597 0.0919 1.0000 3.000 0.5928 0.00831 0.00208 -0.0584 0.0520 1.0000 3.250 0.6138 0.00874 0.00235 -0.0571 0.0113 1.0000 3.500 0.6372 0.00893 0.00258 -0.0561 0.0084 1.0000 3.750 0.6599 0.00920 0.00290 -0.0551 0.0063 1.0000 4.000 0.6828 0.00947 0.00321 -0.0540 0.0050 1.0000 4.250 0.7057 0.00972 0.00347 -0.0530 0.0041 1.0000 4.500 0.7268 0.01018 0.00399 -0.0516 0.0034 1.0000 4.750 0.7488 0.01055 0.00442 -0.0504 0.0031 1.0000 5.000 0.7695 0.01107 0.00504 -0.0488 0.0028 1.0000 5.250 0.7890 0.01170 0.00576 -0.0471 0.0026 1.0000 5.500 0.8083 0.01235 0.00650 -0.0453 0.0025 1.0000 5.750 0.8291 0.01283 0.00702 -0.0440 0.0023 1.0000 6.000 0.8505 0.01324 0.00746 -0.0429 0.0021 1.0000 6.250 0.8652 0.01447 0.00883 -0.0403 0.0019 1.0000 6.500 0.8848 0.01521 0.00967 -0.0386 0.0018 1.0000 6.750 0.9024 0.01637 0.01101 -0.0366 0.0016 1.0000 7.000 0.9187 0.01807 0.01291 -0.0342 0.0015 1.0000 7.250 0.9348 0.02040 0.01551 -0.0316 0.0013 1.0000 7.500 0.9497 0.02293 0.01833 -0.0289 0.0012 1.0000 7.750 0.9640 0.02506 0.02071 -0.0265 0.0011 1.0000 8.000 0.9722 0.02856 0.02459 -0.0229 0.0011 1.0000 8.250 0.9739 0.03308 0.02955 -0.0184 0.0012 1.0000 8.500 0.9738 0.03731 0.03414 -0.0142 0.0012 1.0000 8.750 0.9690 0.04197 0.03912 -0.0099 0.0012 1.0000 9.000 0.9634 0.04606 0.04348 -0.0062 0.0013 1.0000 9.250 0.9550 0.04985 0.04748 -0.0027 0.0014 1.0000 9.500 0.9368 0.05377 0.05158 0.0017 0.0014 1.0000 9.750 0.9207 0.05665 0.05458 0.0050 0.0014 1.0000 10.000 0.8989 0.06075 0.05882 0.0065 0.0014 1.0000 10.250 0.8826 0.06481 0.06301 0.0056 0.0014 1.0000 10.500 0.8624 0.07092 0.06922 0.0016 0.0014 1.0000 10.750 0.8470 0.07808 0.07648 -0.0046 0.0014 1.0000