XFOIL Version 6.94 Calculated polar for: AH-6-40-7 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5184 0.00861 0.00240 -0.1205 0.7786 0.1630 0.500 0.5735 0.00872 0.00231 -0.1195 0.7181 0.1934 1.000 0.6308 0.00871 0.00241 -0.1195 0.6907 0.2536 1.500 0.6869 0.00876 0.00255 -0.1192 0.6421 0.3417 2.500 0.7869 0.00832 0.00297 -0.1161 0.5061 1.0000 3.000 0.8407 0.00900 0.00329 -0.1154 0.4400 1.0000 4.000 0.9485 0.01028 0.00405 -0.1142 0.3493 1.0000 4.500 1.0021 0.01086 0.00444 -0.1137 0.3162 1.0000 5.000 1.0555 0.01146 0.00492 -0.1130 0.2912 1.0000 5.500 1.1092 0.01200 0.00540 -0.1124 0.2709 1.0000 6.000 1.1618 0.01258 0.00593 -0.1117 0.2467 1.0000 6.500 1.2143 0.01316 0.00648 -0.1108 0.2316 1.0000 7.000 1.2662 0.01376 0.00708 -0.1100 0.2090 1.0000 7.500 1.3170 0.01446 0.00775 -0.1089 0.1849 1.0000 8.000 1.3647 0.01545 0.00857 -0.1075 0.1367 1.0000 8.500 1.4001 0.01787 0.01043 -0.1045 0.0603 1.0000 9.000 1.4371 0.01988 0.01232 -0.1014 0.0405 1.0000 9.500 1.4755 0.02147 0.01400 -0.0985 0.0340 1.0000 10.000 1.5090 0.02328 0.01591 -0.0949 0.0305 1.0000 10.500 1.5330 0.02545 0.01820 -0.0901 0.0281 1.0000 11.000 1.5518 0.02746 0.02037 -0.0844 0.0262 1.0000 11.500 1.5568 0.03049 0.02350 -0.0777 0.0250 1.0000 12.000 1.5715 0.03318 0.02645 -0.0731 0.0233 1.0000 12.500 1.5810 0.03657 0.03000 -0.0689 0.0223 1.0000 13.000 1.5877 0.04057 0.03419 -0.0657 0.0207 1.0000 13.500 1.5838 0.04612 0.04000 -0.0627 0.0198 1.0000 14.000 1.5821 0.05202 0.04609 -0.0618 0.0176 1.0000 14.500 1.5812 0.05823 0.05252 -0.0614 0.0178 1.0000 15.000 1.5644 0.06649 0.06100 -0.0607 0.0178 1.0000 15.500 1.5547 0.07565 0.07035 -0.0647 0.0153 1.0000 16.500 1.5248 0.09440 0.08961 -0.0690 0.0152 1.0000 17.000 1.5124 0.10280 0.09820 -0.0697 0.0153 1.0000 17.500 1.4892 0.11655 0.11232 -0.0783 0.0143 1.0000 18.000 1.4724 0.12911 0.12497 -0.0862 0.0126 1.0000 18.500 1.4660 0.13825 0.13434 -0.0900 0.0130 1.0000