XFOIL Version 6.94 Calculated polar for: manu4/26 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1537 0.09875 0.08694 0.0219 0.9999 0.8861 -2.750 -0.1694 0.09621 0.08471 0.0258 0.9999 0.8742 -2.500 -0.1798 0.09353 0.08233 0.0286 0.9999 0.8643 -2.250 -0.2035 0.09083 0.07992 0.0337 0.9999 0.8536 -2.000 -0.2202 0.08804 0.07745 0.0374 0.9999 0.8440 -1.750 -0.2461 0.08521 0.07492 0.0429 0.9999 0.8348 -1.500 -0.2678 0.08243 0.07240 0.0472 0.9999 0.8260 -1.250 -0.2738 0.08010 0.07037 0.0484 0.9999 0.8197 -1.000 -0.2773 0.07834 0.06891 0.0488 0.9999 0.8146 -0.750 -0.2805 0.07777 0.06863 0.0482 0.9999 0.8110 -0.500 -0.2930 0.07937 0.07052 0.0477 0.9999 0.8074 -0.250 -0.3090 0.08182 0.07295 0.0471 0.9999 0.8042 0.000 -0.3126 0.08386 0.07486 0.0449 0.9999 0.8023 0.250 -0.3060 0.08561 0.07645 0.0415 0.9999 0.8019 0.500 -0.2920 0.08722 0.07791 0.0371 0.9999 0.8034 0.750 -0.2732 0.08882 0.07940 0.0321 0.9999 0.8072 1.000 -0.2500 0.09047 0.08098 0.0265 0.9999 0.8139 1.500 -0.1862 0.09409 0.08489 0.0120 0.9999 0.8460 2.000 -0.1198 0.09818 0.08875 -0.0042 0.9999 1.0001 2.250 -0.0985 0.10022 0.09029 -0.0089 0.9999 1.0001 2.500 -0.0734 0.10271 0.09226 -0.0143 0.9999 1.0001 2.750 -0.0456 0.10556 0.09455 -0.0200 0.9999 1.0001 3.000 -0.0153 0.10872 0.09712 -0.0262 0.9999 1.0001 3.250 0.0167 0.11214 0.09992 -0.0325 0.9999 1.0001 3.500 0.0500 0.11581 0.10292 -0.0389 0.9999 1.0001 3.750 0.0851 0.11975 0.10614 -0.0455 0.9999 1.0001 4.000 0.1206 0.12386 0.10949 -0.0519 0.9999 1.0001 4.250 0.1546 0.12803 0.11284 -0.0578 0.9999 1.0001 4.500 0.1846 0.13196 0.11597 -0.0625 0.9999 1.0001 4.750 0.2100 0.13564 0.11891 -0.0661 0.9999 1.0001 5.000 0.2321 0.13907 0.12167 -0.0689 0.9999 1.0001