XFOIL Version 6.94 Calculated polar for: ARA-D BL 12% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0298 0.03596 0.02721 -0.0334 0.9997 0.2649 -2.750 -0.0086 0.03527 0.02659 -0.0327 0.9997 0.2773 -2.500 0.0149 0.03466 0.02595 -0.0324 0.9997 0.2911 -2.250 0.0391 0.03439 0.02569 -0.0321 0.9997 0.3089 -2.000 0.2027 0.02814 0.01869 -0.0412 0.5302 0.4117 -1.750 0.2232 0.02875 0.01834 -0.0383 0.3998 0.4694 -1.500 0.2463 0.02644 0.01741 -0.0343 0.3660 1.0003 -1.250 0.2806 0.02763 0.01743 -0.0344 0.3414 1.0003 -1.000 0.3114 0.02882 0.01793 -0.0342 0.3250 1.0003 -0.750 0.3435 0.02984 0.01867 -0.0342 0.3139 1.0003 -0.250 0.4068 0.03232 0.02064 -0.0342 0.2995 1.0003 0.000 0.4379 0.03359 0.02178 -0.0344 0.2930 1.0003 0.250 0.4680 0.03508 0.02302 -0.0344 0.2864 1.0003 0.500 0.4982 0.03670 0.02457 -0.0347 0.2812 1.0003 0.750 0.5288 0.03821 0.02620 -0.0351 0.2776 1.0003 1.000 0.5590 0.03998 0.02807 -0.0355 0.2756 1.0003 1.250 0.5888 0.04194 0.03016 -0.0361 0.2744 1.0003 1.500 0.6179 0.04411 0.03248 -0.0367 0.2736 1.0003 1.750 0.6458 0.04644 0.03497 -0.0374 0.2723 1.0003 2.000 0.6724 0.04894 0.03763 -0.0381 0.2705 1.0003 2.250 0.6976 0.05173 0.04061 -0.0389 0.2696 1.0003 2.500 0.7207 0.05527 0.04445 -0.0401 0.2722 1.0003 2.750 0.7415 0.05933 0.04877 -0.0416 0.2767 1.0003 3.000 0.7623 0.06337 0.05289 -0.0426 0.2810 1.0003 3.500 0.7878 0.07554 0.06565 -0.0481 0.3052 1.0003 4.000 0.3658 0.11019 0.10273 -0.0637 0.5777 1.0003 4.250 0.3522 0.11102 0.10348 -0.0611 0.5596 1.0003 4.500 0.3650 0.11460 0.10693 -0.0610 0.5484 1.0003 5.000 0.3773 0.11910 0.11124 -0.0592 0.5144 1.0003