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 = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0268 0.04491 0.01945 -0.0221 0.9994 1.0006 -2.750 -0.0062 0.04508 0.01930 -0.0217 0.9994 1.0006 -2.500 0.0142 0.04532 0.01921 -0.0213 0.9994 1.0006 -2.250 0.0344 0.04563 0.01925 -0.0208 0.9994 1.0006 -2.000 0.0544 0.04600 0.01940 -0.0204 0.9994 1.0006 -1.750 0.0744 0.04644 0.01963 -0.0199 0.9994 1.0006 -1.500 0.0941 0.04693 0.01993 -0.0195 0.9994 1.0006 -1.250 0.1137 0.04749 0.02034 -0.0190 0.9994 1.0006 -1.000 0.1331 0.04811 0.02085 -0.0186 0.9994 1.0006 -0.750 0.1523 0.04880 0.02145 -0.0182 0.9994 1.0006 -0.500 0.1714 0.04956 0.02213 -0.0178 0.9994 1.0006 -0.250 0.1904 0.05038 0.02292 -0.0175 0.9994 1.0006 0.000 0.2092 0.05129 0.02381 -0.0173 0.9994 1.0006 0.250 0.2276 0.05227 0.02481 -0.0171 0.9994 1.0006 0.500 0.2458 0.05334 0.02592 -0.0170 0.9994 1.0006 0.750 0.2635 0.05450 0.02717 -0.0170 0.9994 1.0006 1.000 0.2806 0.05578 0.02855 -0.0171 0.9994 1.0006 1.250 0.2970 0.05719 0.03010 -0.0173 0.9994 1.0006 1.500 0.3126 0.05876 0.03184 -0.0176 0.9994 1.0006 1.750 0.3270 0.06053 0.03380 -0.0180 0.9994 1.0006 2.000 0.3398 0.06256 0.03603 -0.0187 0.9994 1.0006 2.250 0.3508 0.06491 0.03860 -0.0196 0.9994 1.0006 2.500 0.3594 0.06764 0.04154 -0.0209 0.9994 1.0006 2.750 0.3658 0.07075 0.04485 -0.0225 0.9994 1.0006 3.000 0.3708 0.07414 0.04838 -0.0243 0.9994 1.0006 3.250 0.3756 0.07767 0.05204 -0.0263 0.9994 1.0006 3.500 0.3809 0.08124 0.05570 -0.0283 0.9994 1.0006 3.750 0.3868 0.08478 0.05933 -0.0303 0.9994 1.0006 4.000 0.3933 0.08829 0.06293 -0.0323 0.9994 1.0006 4.250 0.4003 0.09177 0.06648 -0.0342 0.9994 1.0006 4.500 0.4077 0.09524 0.07003 -0.0361 0.9994 1.0006 4.750 0.4156 0.09870 0.07357 -0.0379 0.9994 1.0006 5.000 0.4238 0.10214 0.07712 -0.0398 0.9994 1.0006