XFOIL Version 6.96 Calculated polar for: GOE 443 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.6465 0.09328 0.08856 0.0065 1.0000 0.0349 -8.000 -0.6485 0.08901 0.08433 0.0030 1.0000 0.0354 -7.750 -0.6501 0.08461 0.07996 -0.0004 1.0000 0.0354 -7.500 -0.6482 0.07999 0.07532 -0.0041 1.0000 0.0358 -7.250 -0.6448 0.07564 0.07079 -0.0092 1.0000 0.0374 -7.000 -0.6390 0.07203 0.06693 -0.0114 1.0000 0.0378 -6.750 -0.6316 0.06843 0.06302 -0.0124 1.0000 0.0380 -6.500 -0.6244 0.06171 0.05644 -0.0127 1.0000 0.0391 -6.250 -0.6145 0.05771 0.05242 -0.0124 1.0000 0.0416 -6.000 -0.6015 0.05376 0.04827 -0.0127 1.0000 0.0435 -5.750 -0.5864 0.04968 0.04392 -0.0128 1.0000 0.0433 -5.250 -0.5430 0.03990 0.03312 -0.0104 1.0000 0.0182 -5.000 -0.5252 0.03615 0.02904 -0.0095 1.0000 0.0174 -4.750 -0.5050 0.03275 0.02522 -0.0083 1.0000 0.0167 -4.500 -0.4831 0.02961 0.02161 -0.0070 1.0000 0.0162 -4.250 -0.4598 0.02670 0.01823 -0.0056 1.0000 0.0160 -4.000 -0.4353 0.02406 0.01512 -0.0043 1.0000 0.0161 -3.750 -0.4100 0.02170 0.01232 -0.0029 1.0000 0.0166 -3.500 -0.3844 0.02007 0.01034 -0.0017 1.0000 0.0188 -3.250 -0.3617 0.01802 0.00822 -0.0005 1.0000 0.0211 -3.000 -0.3396 0.01659 0.00670 0.0010 1.0000 0.0242 -2.750 -0.3177 0.01551 0.00550 0.0023 1.0000 0.0333 -2.500 -0.2957 0.01434 0.00425 0.0037 1.0000 0.0563 -2.250 -0.2866 0.01091 0.00346 0.0066 1.0000 0.6187 -2.000 -0.1885 0.01053 0.00341 -0.0042 1.0000 0.9819 -1.750 -0.1408 0.01039 0.00293 -0.0088 1.0000 1.0000 -1.500 -0.1211 0.01025 0.00257 -0.0076 1.0000 1.0000 -1.250 -0.1011 0.01014 0.00233 -0.0064 1.0000 1.0000 -1.000 -0.0810 0.01006 0.00213 -0.0051 1.0000 1.0000 -0.750 -0.0607 0.01000 0.00199 -0.0039 1.0000 1.0000 -0.500 -0.0404 0.00996 0.00188 -0.0026 1.0000 1.0000 -0.250 -0.0201 0.00994 0.00182 -0.0013 1.0000 1.0000 0.000 0.0000 0.00993 0.00179 0.0000 1.0000 1.0000 0.250 0.0201 0.00994 0.00182 0.0013 1.0000 1.0000 0.500 0.0404 0.00996 0.00188 0.0026 1.0000 1.0000 0.750 0.0607 0.01000 0.00199 0.0039 1.0000 1.0000 1.000 0.0809 0.01006 0.00214 0.0051 1.0000 1.0000 1.250 0.1011 0.01015 0.00233 0.0064 1.0000 1.0000 1.500 0.1210 0.01025 0.00257 0.0076 1.0000 1.0000 1.750 0.1407 0.01039 0.00293 0.0088 1.0000 1.0000 2.000 0.1900 0.01053 0.00343 0.0039 0.9809 1.0000 2.250 0.2866 0.01093 0.00351 -0.0066 0.6128 1.0000 2.500 0.2958 0.01434 0.00425 -0.0037 0.0565 1.0000 2.750 0.3178 0.01553 0.00552 -0.0023 0.0328 1.0000 3.000 0.3398 0.01659 0.00670 -0.0010 0.0242 1.0000 3.250 0.3618 0.01801 0.00821 0.0004 0.0212 1.0000 3.500 0.3845 0.02010 0.01037 0.0017 0.0189 1.0000 3.750 0.4101 0.02172 0.01234 0.0029 0.0166 1.0000 4.000 0.4354 0.02406 0.01513 0.0042 0.0161 1.0000 4.250 0.4599 0.02671 0.01823 0.0056 0.0160 1.0000 4.500 0.4832 0.02959 0.02160 0.0070 0.0162 1.0000 4.750 0.5050 0.03273 0.02520 0.0083 0.0167 1.0000 5.000 0.5251 0.03615 0.02904 0.0095 0.0174 1.0000 5.250 0.5428 0.03988 0.03311 0.0104 0.0182 1.0000 5.500 0.5630 0.04318 0.03677 0.0114 0.0203 1.0000 5.750 0.5479 0.03311 0.02773 0.0134 0.0250 1.0000 6.000 0.5560 0.03838 0.03296 0.0136 0.0284 1.0000 6.500 0.6238 0.06180 0.05646 0.0128 0.0390 1.0000 7.000 0.6391 0.07211 0.06698 0.0116 0.0378 1.0000 7.250 0.6442 0.07552 0.07070 0.0089 0.0372 1.0000 7.500 0.6482 0.07987 0.07519 0.0050 0.0364 1.0000 7.750 0.6500 0.08456 0.07991 0.0009 0.0357 1.0000 8.000 0.6481 0.08900 0.08433 -0.0031 0.0353 1.0000 8.250 0.6462 0.09324 0.08850 -0.0066 0.0345 1.0000