XFOIL Version 6.94 Calculated polar for: nabad 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3805 0.05854 0.04368 -0.0002 0.9999 0.2978 -2.750 -0.3462 0.05556 0.04016 -0.0025 0.9999 0.2951 -2.500 -0.3110 0.05263 0.03683 -0.0042 0.9999 0.2966 -2.250 -0.2749 0.04984 0.03378 -0.0056 0.9999 0.3025 -2.000 -0.2374 0.04737 0.03099 -0.0070 0.9999 0.3130 -1.750 -0.1994 0.04491 0.02843 -0.0083 0.9999 0.3293 -1.500 -0.1617 0.04252 0.02616 -0.0093 0.9999 0.3565 -1.250 -0.1261 0.03984 0.02412 -0.0097 0.9999 0.4036 -0.750 -0.0006 0.03461 0.01924 -0.0170 0.9999 1.0001 -0.500 0.0259 0.03466 0.01868 -0.0157 0.9999 1.0001 -0.250 0.0513 0.03474 0.01845 -0.0145 0.9999 1.0001 0.000 0.0765 0.03484 0.01834 -0.0134 0.9999 1.0001 0.250 0.1014 0.03496 0.01837 -0.0123 0.9999 1.0001 0.500 0.1263 0.03512 0.01854 -0.0114 0.9999 1.0001 0.750 0.1512 0.03532 0.01883 -0.0106 0.9999 1.0001 1.000 0.1757 0.03561 0.01932 -0.0101 0.9999 1.0001 1.250 0.1984 0.03611 0.02016 -0.0099 0.9999 1.0001 1.500 0.2119 0.03740 0.02190 -0.0101 0.9999 1.0001 1.750 0.1943 0.04113 0.02573 -0.0102 0.9999 1.0001 2.000 0.2050 0.04518 0.02950 -0.0163 0.9805 1.0001 2.500 0.3673 0.05035 0.03416 -0.0441 0.8648 1.0001 2.750 0.4028 0.05256 0.03626 -0.0483 0.8448 1.0001 3.000 0.4220 0.05490 0.03847 -0.0497 0.8331 1.0001 3.250 0.4328 0.05728 0.04075 -0.0498 0.8250 1.0001 3.500 0.4457 0.05969 0.04307 -0.0503 0.8199 1.0001 3.750 0.4563 0.06213 0.04545 -0.0504 0.8174 1.0001 4.000 0.4634 0.06459 0.04784 -0.0500 0.8170 1.0001 4.250 0.4678 0.06704 0.05022 -0.0493 0.8185 1.0001 4.500 0.4709 0.06947 0.05259 -0.0485 0.8215 1.0001 4.750 0.4744 0.07188 0.05495 -0.0477 0.8260 1.0001 5.000 0.4808 0.07432 0.05738 -0.0474 0.8317 1.0001