XFOIL Version 6.96 Calculated polar for: FX 71-L-150/25 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.500 -0.8317 0.14602 0.14332 0.0147 1.0000 0.0211 -17.250 -0.8349 0.14073 0.13803 0.0120 1.0000 0.0208 -17.000 -0.9106 0.11729 0.11431 -0.0027 1.0000 0.0201 -16.750 -0.9689 0.10062 0.09730 -0.0135 1.0000 0.0196 -16.500 -1.0772 0.07807 0.07388 -0.0270 1.0000 0.0185 -16.250 -1.0952 0.07217 0.06777 -0.0298 1.0000 0.0185 -16.000 -1.1080 0.06748 0.06287 -0.0319 1.0000 0.0184 -15.750 -1.1169 0.06356 0.05878 -0.0333 1.0000 0.0183 -15.500 -1.1192 0.06037 0.05547 -0.0339 1.0000 0.0182 -15.250 -1.1281 0.05649 0.05141 -0.0346 1.0000 0.0183 -15.000 -1.1285 0.05379 0.04858 -0.0350 1.0000 0.0182 -14.750 -1.1260 0.05119 0.04587 -0.0350 1.0000 0.0182 -14.500 -1.1260 0.04856 0.04309 -0.0349 1.0000 0.0182 -14.250 -1.1227 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0.00415 0.0029 0.2635 0.7415 5.000 0.5446 0.01119 0.00452 0.0035 0.2070 0.7450 5.250 0.5668 0.01182 0.00491 0.0042 0.1549 0.7486 5.500 0.5895 0.01239 0.00529 0.0048 0.1160 0.7527 5.750 0.6123 0.01295 0.00567 0.0054 0.0852 0.7567 6.000 0.6345 0.01349 0.00606 0.0061 0.0570 0.7602 6.250 0.6561 0.01404 0.00651 0.0069 0.0357 0.7639 6.500 0.6784 0.01454 0.00697 0.0077 0.0286 0.7680 6.750 0.7014 0.01497 0.00741 0.0083 0.0256 0.7720 7.000 0.7239 0.01539 0.00783 0.0091 0.0239 0.7755 7.250 0.7464 0.01574 0.00824 0.0098 0.0227 0.7790 7.500 0.7683 0.01614 0.00868 0.0106 0.0221 0.7828 7.750 0.7891 0.01661 0.00917 0.0115 0.0215 0.7869 8.000 0.8082 0.01715 0.00974 0.0127 0.0210 0.7909 8.250 0.8270 0.01761 0.01027 0.0139 0.0208 0.7953 8.500 0.8445 0.01815 0.01086 0.0152 0.0205 0.7996 8.750 0.8606 0.01873 0.01150 0.0167 0.0203 0.8042 9.000 0.8743 0.01932 0.01215 0.0185 0.0201 0.8089 9.250 0.8840 0.01989 0.01280 0.0211 0.0199 0.8145 9.500 0.8928 0.02062 0.01359 0.0234 0.0198 0.8204 9.750 0.9028 0.02138 0.01441 0.0253 0.0195 0.8263 10.000 0.9117 0.02223 0.01535 0.0271 0.0194 0.8337 10.250 0.9213 0.02315 0.01636 0.0286 0.0191 0.8430 10.500 0.9312 0.02402 0.01736 0.0301 0.0190 0.8575 10.750 0.9446 0.02498 0.01854 0.0308 0.0188 0.9020 11.000 0.9843 0.02644 0.02012 0.0259 0.0187 1.0000 11.250 0.9955 0.02770 0.02139 0.0267 0.0186 1.0000 11.500 1.0071 0.02894 0.02265 0.0275 0.0186 1.0000 11.750 1.0183 0.03026 0.02401 0.0283 0.0185 1.0000 12.000 1.0302 0.03154 0.02531 0.0291 0.0184 1.0000 12.250 1.0421 0.03296 0.02675 0.0299 0.0183 1.0000 12.500 1.0546 0.03423 0.02809 0.0306 0.0184 1.0000 12.750 1.0672 0.03571 0.02960 0.0314 0.0183 1.0000 13.000 1.0788 0.03719 0.03114 0.0321 0.0182 1.0000 13.250 1.0895 0.03878 0.03282 0.0327 0.0182 1.0000 13.500 1.1013 0.04034 0.03444 0.0333 0.0183 1.0000 13.750 1.1094 0.04216 0.03636 0.0338 0.0182 1.0000 14.000 1.1164 0.04414 0.03845 0.0342 0.0182 1.0000 14.250 1.1223 0.04624 0.04065 0.0346 0.0182 1.0000 14.500 1.1250 0.04859 0.04313 0.0347 0.0182 1.0000 14.750 1.1295 0.05100 0.04565 0.0349 0.0182 1.0000 15.000 1.1317 0.05347 0.04824 0.0349 0.0183 1.0000 15.250 1.1286 0.05670 0.05162 0.0346 0.0182 1.0000 15.500 1.1254 0.05955 0.05462 0.0339 0.0184 1.0000 15.750 1.1182 0.06345 0.05867 0.0331 0.0183 1.0000 16.000 1.1096 0.06740 0.06278 0.0318 0.0183 1.0000 16.250 1.0970 0.07204 0.06763 0.0297 0.0185 1.0000 16.500 1.0758 0.07841 0.07426 0.0266 0.0186 1.0000 16.750 1.0486 0.08626 0.08239 0.0224 0.0188 1.0000 17.000 0.9101 0.11769 0.11472 0.0022 0.0200 1.0000