XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY MS(1)-0313 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -16.250 -0.9908 0.07431 0.07164 -0.0457 1.0000 0.0116 -16.000 -1.0236 0.06412 0.06122 -0.0526 1.0000 0.0116 -15.750 -1.0427 0.05737 0.05428 -0.0568 1.0000 0.0116 -15.500 -1.0554 0.05213 0.04890 -0.0597 1.0000 0.0116 -15.250 -1.0635 0.04786 0.04450 -0.0618 1.0000 0.0116 -15.000 -1.0687 0.04423 0.04075 -0.0633 1.0000 0.0117 -14.750 -1.0703 0.04117 0.03757 -0.0643 1.0000 0.0118 -14.500 -1.0702 0.03846 0.03475 -0.0649 1.0000 0.0119 -14.250 -1.0677 0.03609 0.03228 -0.0653 1.0000 0.0120 -14.000 -1.0638 0.03395 0.03003 -0.0654 1.0000 0.0121 -13.750 -1.0584 0.03203 0.02801 -0.0653 1.0000 0.0122 -13.500 -1.0520 0.03028 0.02615 -0.0650 1.0000 0.0123 -13.250 -1.0449 0.02871 0.02447 -0.0644 1.0000 0.0125 -13.000 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16.500 1.7913 0.04033 0.03535 -0.0399 0.0258 1.0000 16.750 1.7940 0.04242 0.03753 -0.0388 0.0250 1.0000 17.000 1.7980 0.04443 0.03963 -0.0380 0.0243 1.0000 17.250 1.7995 0.04677 0.04206 -0.0372 0.0236 1.0000 17.500 1.7985 0.04946 0.04483 -0.0365 0.0228 1.0000 17.750 1.7944 0.05259 0.04804 -0.0360 0.0221 1.0000 18.000 1.7869 0.05629 0.05185 -0.0358 0.0213 1.0000 18.250 1.7825 0.05978 0.05545 -0.0359 0.0209 1.0000 18.500 1.7792 0.06325 0.05903 -0.0362 0.0204 1.0000 18.750 1.7720 0.06736 0.06327 -0.0369 0.0199 1.0000 19.000 1.7615 0.07214 0.06816 -0.0381 0.0195 1.0000 19.250 1.7470 0.07770 0.07386 -0.0398 0.0191 1.0000