XFOIL Version 6.96 Calculated polar for: GOE 444 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.6289 0.08588 0.08248 -0.0024 1.0000 0.0157 -8.000 -0.6332 0.08134 0.07799 -0.0061 1.0000 0.0157 -7.750 -0.6377 0.07685 0.07350 -0.0091 1.0000 0.0156 -7.250 -0.6353 0.06739 0.06390 -0.0140 1.0000 0.0162 -6.250 -0.5967 0.05224 0.04779 -0.0145 1.0000 0.0180 -5.250 -0.5525 0.03531 0.02999 -0.0100 1.0000 0.0146 -5.000 -0.5293 0.03178 0.02602 -0.0072 1.0000 0.0087 -4.750 -0.5134 0.02828 0.02214 -0.0053 1.0000 0.0079 -4.500 -0.4943 0.02534 0.01879 -0.0033 1.0000 0.0073 -4.250 -0.4730 0.02265 0.01569 -0.0014 1.0000 0.0069 -4.000 -0.4503 0.02026 0.01291 0.0004 1.0000 0.0066 -3.750 -0.4271 0.01856 0.01092 0.0019 1.0000 0.0073 -3.500 -0.4030 0.01760 0.00974 0.0031 1.0000 0.0090 -3.250 -0.3798 0.01614 0.00805 0.0045 1.0000 0.0093 -3.000 -0.3592 0.01429 0.00610 0.0063 1.0000 0.0098 -2.750 -0.3395 0.01292 0.00462 0.0081 1.0000 0.0111 -2.500 -0.3177 0.01222 0.00381 0.0093 1.0000 0.0153 -2.250 -0.2951 0.01159 0.00297 0.0106 1.0000 0.0201 -2.000 -0.2784 0.01000 0.00230 0.0123 1.0000 0.2455 -1.750 -0.2769 0.00779 0.00227 0.0175 1.0000 0.7408 -1.500 -0.2292 0.00758 0.00239 0.0146 1.0000 0.9058 -1.250 -0.1701 0.00770 0.00230 0.0081 1.0000 0.9549 -1.000 -0.1246 0.00775 0.00221 0.0042 1.0000 0.9734 -0.750 -0.0859 0.00776 0.00213 0.0016 1.0000 0.9858 -0.500 -0.0441 0.00775 0.00206 -0.0017 1.0000 0.9965 -0.250 -0.0169 0.00773 0.00201 -0.0019 1.0000 1.0000 0.000 0.0000 0.00772 0.00198 0.0000 1.0000 1.0000 0.250 0.0169 0.00773 0.00200 0.0019 1.0000 1.0000 0.500 0.0443 0.00775 0.00206 0.0016 0.9964 1.0000 0.750 0.0860 0.00776 0.00214 -0.0017 0.9858 1.0000 1.000 0.1246 0.00775 0.00222 -0.0042 0.9734 1.0000 1.250 0.1701 0.00771 0.00230 -0.0081 0.9547 1.0000 1.500 0.2301 0.00758 0.00239 -0.0148 0.9048 1.0000 1.750 0.2768 0.00787 0.00226 -0.0174 0.7199 1.0000 2.000 0.2792 0.01002 0.00230 -0.0124 0.2415 1.0000 2.250 0.2961 0.01160 0.00298 -0.0108 0.0200 1.0000 2.500 0.3186 0.01225 0.00383 -0.0095 0.0150 1.0000 2.750 0.3404 0.01298 0.00468 -0.0082 0.0109 1.0000 3.000 0.3600 0.01435 0.00617 -0.0064 0.0098 1.0000 3.250 0.3806 0.01616 0.00807 -0.0047 0.0094 1.0000 3.500 0.4040 0.01759 0.00974 -0.0033 0.0090 1.0000 3.750 0.4280 0.01868 0.01106 -0.0020 0.0075 1.0000 4.000 0.4512 0.02030 0.01296 -0.0006 0.0067 1.0000 4.250 0.4737 0.02264 0.01569 0.0012 0.0070 1.0000 4.500 0.4949 0.02532 0.01878 0.0031 0.0074 1.0000 4.750 0.5139 0.02832 0.02218 0.0052 0.0080 1.0000 5.000 0.5295 0.03188 0.02613 0.0071 0.0087 1.0000 5.250 0.5528 0.03532 0.03001 0.0099 0.0146 1.0000 7.250 0.6348 0.06739 0.06390 0.0141 0.0161 1.0000 8.250 0.6283 0.08586 0.08245 0.0025 0.0156 1.0000 8.750 0.6261 0.09454 0.09106 -0.0028 0.0154 1.0000