XFOIL Version 6.94 Calculated polar for: manu02 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 -0.1890 0.06593 0.05785 0.0027 0.3682 0.5865 -2.750 -0.1817 0.06252 0.05444 0.0076 0.3654 0.6356 -2.500 -0.1724 0.05900 0.05095 0.0114 0.3631 0.6755 -2.250 -0.1662 0.05624 0.04815 0.0140 0.3608 0.7025 -2.000 -0.1393 0.05263 0.04451 0.0136 0.3583 0.7261 -1.750 -0.1081 0.05023 0.04197 0.0096 0.3555 0.7267 -1.500 -0.0564 0.04861 0.04007 0.0004 0.3529 0.7103 -1.250 0.2099 0.05143 0.03987 -0.0670 0.3465 0.2679 -1.000 0.2726 0.04946 0.03733 -0.0744 0.3449 0.2562 -0.750 0.3280 0.04832 0.03594 -0.0801 0.3442 0.2596 -0.500 0.3791 0.04751 0.03484 -0.0849 0.3439 0.2636 -0.250 0.4275 0.04688 0.03389 -0.0890 0.3437 0.2666 0.000 0.4862 0.04666 0.03329 -0.0954 0.3426 0.2745 0.250 0.5471 0.04712 0.03332 -0.1024 0.3409 0.2913 0.500 0.6070 0.04727 0.03342 -0.1095 0.3395 0.3292 0.750 0.6817 0.04696 0.03342 -0.1197 0.3389 0.4340 1.000 0.7258 0.04788 0.03458 -0.1235 0.3399 0.5014 1.250 0.7610 0.04814 0.03533 -0.1251 0.3419 0.5570 1.500 0.7883 0.04823 0.03606 -0.1251 0.3458 0.6219 2.000 0.9057 0.05196 0.04092 -0.1404 0.3572 1.0001 2.250 0.9298 0.05407 0.04295 -0.1402 0.3614 1.0001 2.500 0.9546 0.05654 0.04527 -0.1402 0.3651 1.0001 2.750 0.9698 0.05758 0.04655 -0.1383 0.3726 1.0001 3.000 0.9762 0.05972 0.04898 -0.1354 0.3832 1.0001 3.250 0.9940 0.06260 0.05179 -0.1345 0.3904 1.0001 3.500 0.9898 0.06448 0.05408 -0.1304 0.4053 1.0001 3.750 1.0005 0.06800 0.05761 -0.1292 0.4174 1.0001 4.000 0.9684 0.07178 0.06188 -0.1229 0.4418 1.0001 4.250 0.9178 0.07807 0.06861 -0.1169 0.4789 1.0001 4.500 0.3165 0.11611 0.10645 -0.0802 0.8575 0.3743 4.750 0.3408 0.11706 0.10765 -0.0833 0.8401 0.4371 5.000 0.3665 0.11937 0.11024 -0.0862 0.8241 0.4939