XFOIL Version 6.94 Calculated polar for: MH 122 9.32% 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 0.5238 0.00815 0.00351 -0.1377 0.8774 0.6761 0.500 0.5803 0.00826 0.00368 -0.1379 0.8692 0.7059 1.000 0.6313 0.00811 0.00369 -0.1367 0.8593 0.7461 1.500 0.6892 0.00775 0.00350 -0.1366 0.8466 0.7961 2.000 0.7399 0.00742 0.00340 -0.1348 0.8288 0.8626 2.500 0.7920 0.00697 0.00312 -0.1331 0.8090 1.0000 3.000 0.8477 0.00696 0.00317 -0.1329 0.7881 1.0000 3.500 0.9003 0.00693 0.00312 -0.1318 0.7509 1.0000 4.000 0.9456 0.00722 0.00319 -0.1291 0.6641 1.0000 4.500 0.9561 0.00944 0.00408 -0.1200 0.4012 1.0000 5.000 0.9797 0.01159 0.00524 -0.1145 0.2090 1.0000 5.500 1.0139 0.01326 0.00641 -0.1110 0.1002 1.0000 6.000 1.0520 0.01468 0.00758 -0.1080 0.0533 1.0000 6.500 1.0862 0.01640 0.00937 -0.1040 0.0300 1.0000 8.500 1.2275 0.03244 0.02692 -0.0897 0.0058 1.0000 9.000 1.2438 0.03920 0.03440 -0.0838 0.0056 1.0000 9.500 1.2523 0.04316 0.03880 -0.0771 0.0050 1.0000 10.000 1.2332 0.04958 0.04578 -0.0679 0.0049 1.0000 10.500 1.2002 0.05707 0.05378 -0.0598 0.0049 1.0000 11.000 1.1603 0.06593 0.06309 -0.0550 0.0050 1.0000 11.500 1.1167 0.07635 0.07388 -0.0543 0.0051 1.0000 12.000 1.0714 0.09013 0.08798 -0.0598 0.0052 1.0000 13.500 0.9690 0.16098 0.15905 -0.0994 0.0133 1.0000 15.000 0.9664 0.19754 0.19555 -0.1164 0.0157 1.0000 15.500 0.9752 0.20810 0.20610 -0.1213 0.0155 1.0000 16.000 0.9896 0.21697 0.21497 -0.1270 0.0119 1.0000 16.500 0.9980 0.22757 0.22556 -0.1317 0.0103 1.0000 17.000 1.0068 0.23806 0.23606 -0.1361 0.0094 1.0000 17.500 1.0159 0.24833 0.24632 -0.1402 0.0085 1.0000 18.000 1.0247 0.25866 0.25667 -0.1440 0.0079 1.0000 18.500 1.0333 0.26896 0.26698 -0.1476 0.0075 1.0000 19.000 1.0408 0.27959 0.27772 -0.1505 0.0069 1.0000