XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0013 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -1.3474 0.11205 0.10920 0.0169 1.0000 0.0112 -19.500 -1.4000 0.09828 0.09520 0.0094 1.0000 0.0111 -19.250 -1.4357 0.08786 0.08459 0.0037 1.0000 0.0111 -19.000 -1.4619 0.07942 0.07598 -0.0007 1.0000 0.0111 -18.750 -1.4802 0.07263 0.06903 -0.0041 1.0000 0.0112 -18.500 -1.4933 0.06698 0.06324 -0.0066 1.0000 0.0112 -18.250 -1.5028 0.06216 0.05829 -0.0085 1.0000 0.0113 -18.000 -1.5102 0.05784 0.05385 -0.0100 1.0000 0.0114 -17.750 -1.5191 0.05359 0.04949 -0.0111 1.0000 0.0116 -17.500 -1.5248 0.05002 0.04581 -0.0117 1.0000 0.0118 -17.250 -1.5277 0.04690 0.04259 -0.0121 1.0000 0.0120 -17.000 -1.5280 0.04418 0.03978 -0.0121 1.0000 0.0121 -16.750 -1.5264 0.04174 0.03726 -0.0120 1.0000 0.0123 -16.500 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0.8076 8.000 0.9157 0.01310 0.00654 -0.0099 0.0310 0.8121 8.250 0.9428 0.01336 0.00683 -0.0099 0.0299 0.8160 8.500 0.9694 0.01363 0.00714 -0.0099 0.0285 0.8201 8.750 0.9957 0.01395 0.00746 -0.0099 0.0271 0.8246 9.000 1.0215 0.01432 0.00784 -0.0098 0.0257 0.8292 9.250 1.0475 0.01463 0.00820 -0.0097 0.0250 0.8337 9.500 1.0731 0.01497 0.00859 -0.0095 0.0241 0.8386 9.750 1.0983 0.01533 0.00898 -0.0094 0.0232 0.8435 10.000 1.1232 0.01573 0.00940 -0.0092 0.0223 0.8480 10.250 1.1472 0.01619 0.00989 -0.0088 0.0213 0.8532 10.500 1.1715 0.01657 0.01034 -0.0086 0.0209 0.8590 10.750 1.1955 0.01699 0.01081 -0.0082 0.0204 0.8646 11.000 1.2187 0.01741 0.01130 -0.0078 0.0198 0.8707 11.250 1.2417 0.01787 0.01180 -0.0073 0.0192 0.8771 11.500 1.2637 0.01835 0.01233 -0.0068 0.0186 0.8840 11.750 1.2848 0.01890 0.01294 -0.0061 0.0181 0.8926 12.000 1.3042 0.01953 0.01364 -0.0051 0.0175 0.9021 12.250 1.3244 0.02002 0.01421 -0.0043 0.0173 0.9130 12.500 1.3432 0.02052 0.01481 -0.0032 0.0169 0.9270 12.750 1.3597 0.02101 0.01542 -0.0017 0.0166 0.9523 13.000 1.3777 0.02160 0.01608 -0.0008 0.0162 1.0000 13.250 1.3923 0.02229 0.01683 0.0006 0.0159 1.0000 13.500 1.4061 0.02309 0.01767 0.0020 0.0156 1.0000 13.750 1.4198 0.02397 0.01861 0.0031 0.0153 1.0000 14.000 1.4327 0.02495 0.01963 0.0042 0.0149 1.0000 14.250 1.4445 0.02606 0.02080 0.0052 0.0146 1.0000 14.500 1.4545 0.02734 0.02215 0.0063 0.0142 1.0000 14.750 1.4672 0.02843 0.02331 0.0071 0.0140 1.0000 15.000 1.4785 0.02964 0.02460 0.0079 0.0138 1.0000 15.250 1.4889 0.03096 0.02600 0.0086 0.0136 1.0000 15.500 1.4982 0.03240 0.02751 0.0094 0.0134 1.0000 15.750 1.5066 0.03395 0.02915 0.0100 0.0132 1.0000 16.000 1.5139 0.03564 0.03092 0.0106 0.0130 1.0000 16.250 1.5199 0.03749 0.03285 0.0111 0.0127 1.0000 16.500 1.5248 0.03948 0.03492 0.0115 0.0125 1.0000 16.750 1.5282 0.04169 0.03721 0.0117 0.0123 1.0000 17.000 1.5299 0.04412 0.03972 0.0119 0.0121 1.0000 17.250 1.5298 0.04683 0.04252 0.0118 0.0120 1.0000 17.500 1.5271 0.04994 0.04573 0.0115 0.0118 1.0000 17.750 1.5212 0.05352 0.04941 0.0109 0.0116 1.0000 18.000 1.5122 0.05779 0.05380 0.0097 0.0114 1.0000 18.250 1.5057 0.06198 0.05811 0.0083 0.0113 1.0000 18.500 1.4961 0.06681 0.06308 0.0064 0.0112 1.0000 18.750 1.4826 0.07253 0.06893 0.0038 0.0112 1.0000 19.000 1.4648 0.07925 0.07581 0.0004 0.0111 1.0000 19.250 1.4392 0.08758 0.08430 -0.0040 0.0111 1.0000