XFOIL Version 6.94 Calculated polar for: Eif385mod3 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 0.0957 0.03397 0.02259 -0.0404 0.3728 0.4989 -2.750 0.1301 0.03440 0.02222 -0.0420 0.3565 0.4819 -2.500 0.1644 0.03437 0.02186 -0.0429 0.3445 0.4692 -2.250 0.1990 0.03484 0.02180 -0.0439 0.3340 0.4582 -2.000 0.2336 0.03514 0.02173 -0.0448 0.3265 0.4535 -1.750 0.2695 0.03523 0.02169 -0.0457 0.3218 0.4480 -1.500 0.3061 0.03552 0.02181 -0.0467 0.3182 0.4413 -1.250 0.3428 0.03579 0.02199 -0.0477 0.3151 0.4400 -1.000 0.3795 0.03611 0.02227 -0.0488 0.3126 0.4400 -0.750 0.4148 0.03648 0.02266 -0.0497 0.3108 0.4424 -0.500 0.4488 0.03700 0.02321 -0.0505 0.3094 0.4439 -0.250 0.4818 0.03764 0.02391 -0.0511 0.3086 0.4448 0.000 0.5136 0.03833 0.02472 -0.0516 0.3082 0.4494 0.250 0.5449 0.03915 0.02568 -0.0521 0.3083 0.4530 0.500 0.5755 0.04009 0.02680 -0.0525 0.3086 0.4560 1.000 0.6380 0.04090 0.02922 -0.0535 0.3078 1.0007 1.250 0.6666 0.04238 0.03072 -0.0538 0.3072 1.0007 1.500 0.6945 0.04394 0.03233 -0.0540 0.3069 1.0007 1.750 0.7216 0.04558 0.03411 -0.0543 0.3078 1.0007 2.000 0.7478 0.04743 0.03618 -0.0546 0.3109 1.0007 2.250 0.7729 0.04955 0.03851 -0.0549 0.3147 1.0007 2.500 0.7968 0.05191 0.04104 -0.0553 0.3186 1.0007 2.750 0.8202 0.05452 0.04375 -0.0556 0.3225 1.0007 3.000 0.8436 0.05750 0.04674 -0.0559 0.3261 1.0007 3.250 0.8572 0.06010 0.05001 -0.0564 0.3369 1.0007 3.500 0.8666 0.06457 0.05482 -0.0570 0.3486 1.0007 3.750 0.8726 0.06919 0.05979 -0.0579 0.3631 1.0007 4.000 0.8744 0.07581 0.06661 -0.0595 0.3826 1.0007 4.250 0.4613 0.11545 0.10755 -0.0738 0.7385 1.0007 4.500 0.4738 0.11821 0.11021 -0.0737 0.7257 1.0007 4.750 0.5088 0.12309 0.11498 -0.0769 0.7164 1.0007 5.000 0.5207 0.12542 0.11722 -0.0764 0.6996 1.0007