XFOIL Version 6.94 Calculated polar for: AH 82-150 A 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 0.3826 0.01546 0.01010 -0.0839 0.7388 0.8264 0.500 0.4431 0.01523 0.00977 -0.0858 0.7313 0.8301 1.000 0.5010 0.01488 0.00932 -0.0867 0.7253 0.8328 1.500 0.5547 0.01490 0.00933 -0.0869 0.7180 0.8358 2.000 0.6085 0.01481 0.00928 -0.0875 0.7097 0.8380 2.500 0.6672 0.01462 0.00913 -0.0887 0.7021 0.8402 3.000 0.7294 0.01439 0.00886 -0.0905 0.6952 0.8424 3.500 0.7842 0.01451 0.00905 -0.0912 0.6868 0.8460 4.000 0.8404 0.01440 0.00904 -0.0921 0.6767 0.8485 4.500 0.9027 0.01411 0.00883 -0.0938 0.6674 0.8507 5.000 0.9615 0.01414 0.00894 -0.0951 0.6572 0.8528 5.500 1.0151 0.01391 0.00889 -0.0953 0.6436 0.8552 6.000 1.0683 0.01314 0.00816 -0.0945 0.6168 0.8581 6.500 1.1127 0.01273 0.00764 -0.0921 0.5642 0.8613 7.000 1.1422 0.01334 0.00797 -0.0877 0.4813 0.8655 7.500 1.0796 0.01810 0.01113 -0.0704 0.2384 0.8741 8.000 1.0536 0.02302 0.01510 -0.0621 0.0745 0.8798 8.500 1.0516 0.02697 0.01886 -0.0567 0.0207 0.8857 9.000 1.0753 0.02933 0.02144 -0.0544 0.0200 0.8917 9.500 1.0862 0.03280 0.02517 -0.0513 0.0185 0.8984 10.000 1.0878 0.03685 0.02948 -0.0473 0.0176 0.9062 10.500 1.0925 0.04107 0.03391 -0.0438 0.0169 0.9152 11.000 1.1143 0.04347 0.03655 -0.0419 0.0153 0.9307 11.500 1.1239 0.04655 0.03988 -0.0378 0.0158 1.0000 12.000 1.1501 0.04979 0.04327 -0.0370 0.0143 1.0000 12.500 1.1773 0.05314 0.04682 -0.0357 0.0133 1.0000 13.000 1.1943 0.05739 0.05118 -0.0350 0.0112 1.0000 13.500 1.2207 0.06268 0.05684 -0.0323 0.0101 1.0000 14.000 1.2266 0.06926 0.06384 -0.0308 0.0101 1.0000 14.500 1.2214 0.07576 0.07074 -0.0301 0.0105 1.0000 15.500 1.1328 0.10382 0.10032 -0.0336 0.0148 1.0000 16.000 1.0893 0.11873 0.11560 -0.0405 0.0157 1.0000