XFOIL Version 6.94 Calculated polar for: NACA M15 AIRFOIL 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 0.3154 0.00976 0.00274 -0.0503 0.6321 0.2629 1.500 0.5718 0.00853 0.00328 -0.0684 0.5892 1.0000 2.000 0.6241 0.00855 0.00322 -0.0679 0.5723 1.0000 2.500 0.6759 0.00862 0.00321 -0.0672 0.5535 1.0000 3.000 0.7268 0.00873 0.00323 -0.0664 0.5323 1.0000 3.500 0.7736 0.00899 0.00317 -0.0648 0.4767 1.0000 4.000 0.8191 0.00950 0.00337 -0.0632 0.4234 1.0000 4.500 0.8652 0.00995 0.00365 -0.0616 0.3914 1.0000 5.000 0.9092 0.01052 0.00399 -0.0597 0.3442 1.0000 5.500 0.9506 0.01126 0.00443 -0.0573 0.2860 1.0000 6.000 0.9776 0.01308 0.00554 -0.0529 0.1629 1.0000 6.500 0.9911 0.01567 0.00733 -0.0464 0.0239 1.0000 7.000 1.0259 0.01667 0.00827 -0.0430 0.0042 1.0000 7.500 1.0640 0.01746 0.00913 -0.0403 0.0039 1.0000 8.000 1.1000 0.01838 0.01015 -0.0374 0.0038 1.0000 8.500 1.1303 0.01945 0.01132 -0.0336 0.0038 1.0000 9.000 1.1559 0.02075 0.01275 -0.0294 0.0038 1.0000 9.500 1.1807 0.02239 0.01453 -0.0259 0.0039 1.0000 10.000 1.2035 0.02443 0.01672 -0.0229 0.0040 1.0000 10.500 1.2234 0.02696 0.01942 -0.0203 0.0041 1.0000 11.000 1.2394 0.03008 0.02272 -0.0181 0.0042 1.0000 11.500 1.2503 0.03395 0.02678 -0.0163 0.0044 1.0000 12.000 1.2559 0.03858 0.03160 -0.0149 0.0045 1.0000 12.500 1.2555 0.04404 0.03725 -0.0139 0.0046 1.0000 13.000 1.2499 0.05026 0.04365 -0.0133 0.0047 1.0000 13.500 1.2411 0.05714 0.05071 -0.0131 0.0048 1.0000 14.000 1.2305 0.06450 0.05821 -0.0133 0.0049 1.0000 14.500 1.2311 0.07083 0.06470 -0.0138 0.0050 1.0000 15.000 1.2355 0.07700 0.07104 -0.0148 0.0052 1.0000 15.500 1.2375 0.08350 0.07774 -0.0157 0.0055 1.0000 16.000 1.2377 0.09002 0.08446 -0.0163 0.0059 1.0000 16.500 1.2391 0.09613 0.09072 -0.0165 0.0063 1.0000 17.000 1.2528 0.09933 0.09395 -0.0142 0.0067 1.0000 17.500 1.2489 0.10800 0.10294 -0.0180 0.0070 1.0000 18.000 1.2412 0.11644 0.11184 -0.0198 0.0079 1.0000