XFOIL Version 6.94 Calculated polar for: HQ 3.5/10 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 0.4798 0.00707 0.00239 -0.1047 0.7827 0.7756 0.500 0.5343 0.00705 0.00229 -0.1042 0.7646 0.7956 1.000 0.5883 0.00703 0.00227 -0.1037 0.7465 0.8175 1.500 0.6414 0.00701 0.00228 -0.1029 0.7270 0.8441 2.000 0.6920 0.00692 0.00228 -0.1015 0.7052 0.8859 2.500 0.7537 0.00688 0.00230 -0.1027 0.6797 1.0000 3.000 0.8081 0.00713 0.00242 -0.1025 0.6491 1.0000 3.500 0.8604 0.00744 0.00262 -0.1019 0.6079 1.0000 4.000 0.9023 0.00833 0.00292 -0.0993 0.4892 1.0000 4.500 0.9402 0.00976 0.00362 -0.0964 0.3509 1.0000 5.000 0.9818 0.01103 0.00433 -0.0944 0.2586 1.0000 5.500 1.0294 0.01179 0.00492 -0.0933 0.2214 1.0000 6.000 1.0726 0.01290 0.00561 -0.0916 0.1477 1.0000 6.500 1.1094 0.01456 0.00680 -0.0890 0.0654 1.0000 7.000 1.1427 0.01654 0.00843 -0.0856 0.0035 1.0000 7.500 1.1862 0.01746 0.00950 -0.0837 0.0031 1.0000 8.000 1.2263 0.01861 0.01085 -0.0812 0.0029 1.0000 8.500 1.2607 0.02002 0.01254 -0.0778 0.0029 1.0000 9.000 1.2862 0.02186 0.01465 -0.0731 0.0029 1.0000 9.500 1.3036 0.02419 0.01726 -0.0676 0.0030 1.0000 10.000 1.3119 0.02729 0.02065 -0.0617 0.0031 1.0000 10.500 1.3149 0.03117 0.02483 -0.0561 0.0032 1.0000 11.000 1.3156 0.03584 0.02982 -0.0513 0.0034 1.0000 11.500 1.3135 0.04161 0.03596 -0.0472 0.0035 1.0000 12.000 1.3055 0.04926 0.04404 -0.0438 0.0038 1.0000 12.500 1.3088 0.05425 0.04933 -0.0423 0.0039 1.0000 13.000 1.3069 0.06033 0.05572 -0.0418 0.0041 1.0000 13.500 1.2960 0.06839 0.06418 -0.0424 0.0045 1.0000 14.000 1.2584 0.08238 0.07883 -0.0454 0.0053 1.0000 14.500 1.2249 0.09665 0.09352 -0.0521 0.0055 1.0000 15.000 1.1883 0.11365 0.11089 -0.0620 0.0057 1.0000 15.500 1.1600 0.13061 0.12813 -0.0730 0.0057 1.0000 16.000 1.1329 0.14880 0.14656 -0.0851 0.0056 1.0000 16.500 1.1111 0.16690 0.16479 -0.0969 0.0054 1.0000