XFOIL Version 6.94 Calculated polar for: GOA 2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1671 0.02990 0.01604 -0.0324 1.0000 0.2465 -2.750 -0.1400 0.02720 0.01389 -0.0310 1.0000 0.3093 -2.500 -0.1125 0.02209 0.01115 -0.0266 1.0000 1.0000 -2.250 -0.0863 0.02208 0.01017 -0.0252 1.0000 1.0000 -2.000 -0.0616 0.02208 0.00956 -0.0240 1.0000 1.0000 -1.750 -0.0372 0.02211 0.00912 -0.0230 1.0000 1.0000 -1.500 -0.0130 0.02214 0.00877 -0.0220 1.0000 1.0000 -1.250 0.0112 0.02219 0.00850 -0.0210 1.0000 1.0000 -1.000 0.0354 0.02226 0.00834 -0.0201 1.0000 1.0000 -0.750 0.0596 0.02234 0.00829 -0.0192 1.0000 1.0000 -0.500 0.0841 0.02244 0.00834 -0.0182 1.0000 1.0000 -0.250 0.1092 0.02254 0.00858 -0.0172 1.0000 1.0000 0.000 0.1358 0.02262 0.00905 -0.0161 1.0000 1.0000 0.250 0.1673 0.02260 0.00977 -0.0152 1.0000 1.0000 0.500 0.2579 0.02639 0.01086 -0.0228 0.3828 1.0000 0.750 0.2843 0.02794 0.01188 -0.0218 0.3583 1.0000 1.000 0.3141 0.02941 0.01307 -0.0214 0.3425 1.0000 1.250 0.3455 0.03085 0.01441 -0.0213 0.3284 1.0000 1.500 0.3769 0.03236 0.01589 -0.0213 0.3134 1.0000 1.750 0.4087 0.03393 0.01755 -0.0216 0.3010 1.0000 2.000 0.4415 0.03564 0.01942 -0.0220 0.2956 1.0000 2.250 0.4753 0.03742 0.02160 -0.0227 0.2959 1.0000 2.500 0.5088 0.03934 0.02398 -0.0237 0.2984 1.0000 2.750 0.5414 0.04148 0.02661 -0.0248 0.3028 1.0000 3.000 0.5721 0.04396 0.02947 -0.0257 0.3078 1.0000 3.250 0.6041 0.04647 0.03264 -0.0279 0.3172 1.0000 3.500 0.6338 0.04960 0.03611 -0.0294 0.3260 1.0000 3.750 0.6641 0.05284 0.03995 -0.0327 0.3412 1.0000 4.000 0.6930 0.05656 0.04412 -0.0363 0.3591 1.0000 4.250 0.7186 0.06074 0.04892 -0.0431 0.3865 1.0000 4.500 0.7389 0.06556 0.05415 -0.0508 0.4194 1.0000 4.750 0.7553 0.07101 0.05984 -0.0594 0.4604 1.0000 5.000 0.7512 0.07687 0.06582 -0.0706 0.5136 1.0000