XFOIL Version 6.94 Calculated polar for: nacc632 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 -0.0148 0.02375 0.01297 -0.0251 0.9994 1.0006 -2.750 0.0067 0.02429 0.01331 -0.0248 0.9994 1.0006 -2.500 0.0280 0.02490 0.01372 -0.0245 0.9994 1.0006 -2.250 0.0495 0.02557 0.01422 -0.0243 0.9994 1.0006 -2.000 0.0710 0.02628 0.01479 -0.0241 0.9994 1.0006 -1.750 0.0926 0.02703 0.01542 -0.0240 0.9994 1.0006 -1.500 0.1140 0.02782 0.01610 -0.0239 0.9994 1.0006 -1.250 0.1354 0.02865 0.01685 -0.0239 0.9994 1.0006 -1.000 0.1567 0.02952 0.01766 -0.0239 0.9994 1.0006 -0.750 0.1778 0.03043 0.01853 -0.0240 0.9994 1.0006 -0.500 0.1986 0.03139 0.01946 -0.0241 0.9994 1.0006 -0.250 0.2193 0.03240 0.02047 -0.0242 0.9994 1.0006 0.000 0.2396 0.03347 0.02156 -0.0244 0.9994 1.0006 0.250 0.2595 0.03462 0.02274 -0.0247 0.9994 1.0006 0.500 0.2789 0.03587 0.02405 -0.0251 0.9994 1.0006 0.750 0.3274 0.03643 0.02473 -0.0309 0.9866 1.0006 1.000 0.4495 0.03447 0.02304 -0.0484 0.9390 1.0006 1.250 0.5666 0.03105 0.01998 -0.0611 0.8761 1.0006 1.500 0.6066 0.02968 0.01871 -0.0574 0.8019 1.0006 1.750 0.6219 0.02903 0.01791 -0.0494 0.7164 1.0006 2.000 0.6383 0.02887 0.01732 -0.0434 0.6293 1.0006 2.250 0.6600 0.02925 0.01722 -0.0400 0.5618 1.0006 2.500 0.6845 0.03000 0.01755 -0.0382 0.5114 1.0006 2.750 0.7112 0.03103 0.01833 -0.0374 0.4739 1.0006 3.000 0.7383 0.03220 0.01925 -0.0369 0.4438 1.0006 3.250 0.7659 0.03348 0.02031 -0.0366 0.4176 1.0006 3.500 0.7940 0.03500 0.02171 -0.0365 0.3961 1.0006 3.750 0.8218 0.03663 0.02338 -0.0367 0.3758 1.0006 4.000 0.8497 0.03844 0.02517 -0.0368 0.3585 1.0006 4.250 0.8780 0.04056 0.02737 -0.0373 0.3457 1.0006 4.500 0.9060 0.04280 0.02964 -0.0376 0.3345 1.0006 4.750 0.9326 0.04528 0.03256 -0.0384 0.3225 1.0006 5.000 0.9579 0.04780 0.03529 -0.0389 0.3103 1.0006