XFOIL Version 6.96 Calculated polar for: RAF 30 MOD AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.000 -0.6494 0.09513 0.09023 0.0046 1.0000 0.1321 -8.750 -0.6676 0.09049 0.08573 -0.0009 1.0000 0.1370 -8.500 -0.6934 0.08525 0.08056 -0.0072 1.0000 0.1378 -8.250 -0.6677 0.08225 0.07754 -0.0025 1.0000 0.1474 -8.000 -0.7081 0.07691 0.07215 -0.0107 1.0000 0.1519 -7.750 -0.6814 0.07341 0.06873 -0.0066 1.0000 0.1596 -7.500 -0.6924 0.06866 0.06392 -0.0093 1.0000 0.1692 -7.250 -0.6968 0.06477 0.05991 -0.0105 1.0000 0.1817 -7.000 -0.6916 0.06128 0.05635 -0.0103 1.0000 0.1955 -6.750 -0.6911 0.04454 0.03764 -0.0142 1.0000 0.0833 -6.500 -0.6755 0.03838 0.03085 -0.0125 1.0000 0.0697 -6.250 -0.6578 0.03426 0.02639 -0.0112 1.0000 0.0668 -6.000 -0.6382 0.03095 0.02255 -0.0097 1.0000 0.0657 -5.750 -0.6168 0.02866 0.01976 -0.0082 1.0000 0.0686 -5.500 -0.5934 0.02629 0.01695 -0.0068 1.0000 0.0697 -5.250 -0.5686 0.02429 0.01458 -0.0055 1.0000 0.0712 -5.000 -0.5450 0.02208 0.01244 -0.0048 1.0000 0.0772 -4.750 -0.5201 0.02078 0.01094 -0.0037 1.0000 0.0834 -4.500 -0.4971 0.01906 0.00934 -0.0026 1.0000 0.0897 -4.250 -0.4752 0.01791 0.00823 -0.0013 1.0000 0.1014 -4.000 -0.4546 0.01675 0.00714 0.0003 1.0000 0.1145 -3.750 -0.4358 0.01543 0.00607 0.0020 1.0000 0.1464 -3.500 -0.4332 0.01241 0.00531 0.0064 1.0000 0.5186 -3.250 -0.4183 0.01188 0.00528 0.0101 1.0000 0.6627 -3.000 -0.3998 0.01168 0.00521 0.0132 1.0000 0.7405 -2.750 -0.3810 0.01162 0.00528 0.0167 1.0000 0.8128 -2.500 -0.3528 0.01183 0.00553 0.0190 1.0000 0.8775 -2.250 -0.2938 0.01248 0.00595 0.0154 1.0000 0.9304 -2.000 -0.2113 0.01289 0.00600 0.0056 1.0000 0.9581 -1.750 -0.1461 0.01275 0.00564 -0.0021 1.0000 0.9756 -1.500 -0.0834 0.01244 0.00515 -0.0098 1.0000 0.9926 -1.250 -0.0466 0.01206 0.00468 -0.0128 1.0000 1.0000 -1.000 -0.0353 0.01175 0.00437 -0.0109 1.0000 1.0000 -0.750 -0.0254 0.01150 0.00412 -0.0086 1.0000 1.0000 -0.500 -0.0167 0.01131 0.00393 -0.0058 1.0000 1.0000 -0.250 -0.0084 0.01119 0.00382 -0.0029 1.0000 1.0000 0.000 0.0000 0.01116 0.00379 0.0000 1.0000 1.0000 0.250 0.0084 0.01119 0.00382 0.0029 1.0000 1.0000 0.500 0.0167 0.01131 0.00393 0.0058 1.0000 1.0000 0.750 0.0254 0.01150 0.00412 0.0086 1.0000 1.0000 1.000 0.0353 0.01175 0.00437 0.0109 1.0000 1.0000 1.250 0.0466 0.01206 0.00468 0.0128 1.0000 1.0000 1.500 0.0833 0.01244 0.00515 0.0098 0.9926 1.0000 1.750 0.1460 0.01275 0.00564 0.0021 0.9756 1.0000 2.000 0.2115 0.01289 0.00600 -0.0056 0.9581 1.0000 2.250 0.2938 0.01248 0.00595 -0.0154 0.9304 1.0000 2.500 0.3529 0.01184 0.00554 -0.0191 0.8780 1.0000 2.750 0.3810 0.01162 0.00528 -0.0167 0.8122 1.0000 3.000 0.4000 0.01168 0.00521 -0.0133 0.7403 1.0000 3.250 0.4184 0.01188 0.00528 -0.0102 0.6623 1.0000 3.500 0.4333 0.01242 0.00531 -0.0064 0.5170 1.0000 3.750 0.4359 0.01544 0.00608 -0.0021 0.1458 1.0000 4.000 0.4547 0.01675 0.00715 -0.0003 0.1145 1.0000 4.250 0.4753 0.01791 0.00823 0.0013 0.1013 1.0000 4.500 0.4972 0.01906 0.00934 0.0026 0.0896 1.0000 4.750 0.5202 0.02078 0.01094 0.0037 0.0834 1.0000 5.000 0.5451 0.02209 0.01246 0.0048 0.0774 1.0000 5.250 0.5687 0.02430 0.01459 0.0055 0.0712 1.0000 5.500 0.5934 0.02629 0.01695 0.0068 0.0697 1.0000 5.750 0.6169 0.02868 0.01978 0.0082 0.0688 1.0000 6.000 0.6382 0.03094 0.02254 0.0097 0.0656 1.0000 6.250 0.6578 0.03425 0.02639 0.0112 0.0668 1.0000 6.500 0.6754 0.03838 0.03085 0.0125 0.0697 1.0000 6.750 0.6912 0.04463 0.03772 0.0142 0.0834 1.0000 8.500 0.6964 0.08517 0.08047 0.0077 0.1379 1.0000 8.750 0.6660 0.09053 0.08576 0.0005 0.1368 1.0000 9.000 0.6497 0.09510 0.09024 -0.0046 0.1319 1.0000