XFOIL Version 6.94 Calculated polar for: manu4/36 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4259 0.06407 0.05713 0.0623 0.9999 0.7963 -2.750 -0.4750 0.06183 0.05499 0.0681 0.9999 0.7620 -2.500 -0.4541 0.06101 0.05399 0.0549 0.9999 0.6749 -1.750 -0.2633 0.05759 0.04950 0.0083 0.9999 0.4023 -1.250 -0.0494 0.04722 0.03850 -0.0229 0.7638 0.3537 -1.000 0.0938 0.04323 0.03135 -0.0422 0.5173 0.3511 -0.750 0.1455 0.04265 0.02988 -0.0470 0.4974 0.3627 -0.500 0.2090 0.04184 0.02837 -0.0541 0.4827 0.3787 -0.250 0.2830 0.04109 0.02690 -0.0633 0.4712 0.4051 0.000 0.3668 0.03964 0.02532 -0.0745 0.4617 0.4863 0.250 0.5289 0.03776 0.02419 -0.1026 0.4455 1.0001 0.500 0.5764 0.03871 0.02467 -0.1068 0.4395 1.0001 0.750 0.6207 0.03969 0.02526 -0.1102 0.4353 1.0001 1.000 0.6631 0.04071 0.02600 -0.1132 0.4329 1.0001 1.250 0.7029 0.04176 0.02687 -0.1158 0.4318 1.0001 1.500 0.7395 0.04284 0.02786 -0.1178 0.4317 1.0001 1.750 0.7733 0.04400 0.02900 -0.1192 0.4319 1.0001 2.000 0.8039 0.04522 0.03025 -0.1202 0.4328 1.0001 2.250 0.8320 0.04653 0.03162 -0.1207 0.4342 1.0001 2.500 0.8581 0.04798 0.03314 -0.1209 0.4359 1.0001 2.750 0.8824 0.04954 0.03477 -0.1208 0.4381 1.0001 3.000 0.9050 0.05125 0.03659 -0.1206 0.4406 1.0001 3.250 0.9269 0.05311 0.03852 -0.1202 0.4432 1.0001 3.500 0.9497 0.05518 0.04062 -0.1201 0.4458 1.0001 3.750 0.9553 0.05697 0.04276 -0.1176 0.4506 1.0001 4.000 0.9587 0.05940 0.04545 -0.1151 0.4559 1.0001 4.250 0.9642 0.06212 0.04833 -0.1131 0.4608 1.0001 4.500 0.9733 0.06500 0.05133 -0.1117 0.4649 1.0001 4.750 0.9889 0.06792 0.05429 -0.1112 0.4684 1.0001 5.000 0.9203 0.07378 0.06052 -0.1024 0.4818 1.0001