XFOIL Version 6.96 Calculated polar for: FX 71-L-150/20 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.250 -1.0086 0.14203 0.13861 0.0206 1.0000 0.0116 -19.000 -1.0264 0.13340 0.12982 0.0152 1.0000 0.0115 -18.750 -1.0411 0.12575 0.12201 0.0106 1.0000 0.0115 -18.500 -1.0534 0.11879 0.11492 0.0064 1.0000 0.0115 -18.250 -1.0658 0.11200 0.10798 0.0024 1.0000 0.0115 -18.000 -1.0762 0.10580 0.10165 -0.0011 1.0000 0.0115 -17.750 -1.0856 0.09998 0.09570 -0.0043 1.0000 0.0115 -17.500 -1.0933 0.09457 0.09016 -0.0072 1.0000 0.0115 -17.250 -1.1001 0.08949 0.08495 -0.0097 1.0000 0.0114 -17.000 -1.1058 0.08473 0.08007 -0.0121 1.0000 0.0115 -16.750 -1.1101 0.08030 0.07554 -0.0141 1.0000 0.0115 -16.500 -1.1139 0.07610 0.07122 -0.0159 1.0000 0.0115 -16.250 -1.1162 0.07222 0.06724 -0.0176 1.0000 0.0115 -16.000 -1.1173 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0.6445 3.500 0.3868 0.01045 0.00314 0.0009 0.2045 0.6479 3.750 0.4113 0.01098 0.00346 0.0012 0.1506 0.6513 4.000 0.4364 0.01145 0.00375 0.0015 0.1100 0.6546 4.250 0.4613 0.01191 0.00406 0.0017 0.0769 0.6580 4.500 0.4869 0.01229 0.00434 0.0020 0.0550 0.6614 4.750 0.5130 0.01262 0.00460 0.0022 0.0395 0.6646 5.250 0.5647 0.01323 0.00516 0.0026 0.0257 0.6701 5.500 0.5903 0.01353 0.00545 0.0029 0.0175 0.6731 5.750 0.6157 0.01384 0.00575 0.0032 0.0158 0.6762 6.000 0.6412 0.01414 0.00607 0.0035 0.0150 0.6795 6.250 0.6664 0.01447 0.00641 0.0038 0.0145 0.6831 6.500 0.6913 0.01480 0.00676 0.0042 0.0141 0.6869 6.750 0.7158 0.01514 0.00714 0.0046 0.0138 0.6906 7.000 0.7395 0.01552 0.00757 0.0051 0.0135 0.6944 7.250 0.7630 0.01591 0.00800 0.0057 0.0133 0.6980 7.500 0.7856 0.01636 0.00849 0.0063 0.0131 0.7017 7.750 0.8074 0.01687 0.00902 0.0070 0.0129 0.7053 8.000 0.8294 0.01731 0.00951 0.0077 0.0128 0.7089 8.250 0.8501 0.01779 0.01005 0.0085 0.0127 0.7128 8.500 0.8699 0.01830 0.01062 0.0094 0.0127 0.7171 8.750 0.8890 0.01884 0.01121 0.0104 0.0126 0.7214 9.000 0.9064 0.01944 0.01185 0.0116 0.0125 0.7255 9.250 0.9223 0.02006 0.01253 0.0130 0.0125 0.7295 9.500 0.9336 0.02069 0.01323 0.0150 0.0124 0.7338 9.750 0.9428 0.02142 0.01403 0.0172 0.0124 0.7386 10.000 0.9503 0.02238 0.01506 0.0190 0.0123 0.7437 10.500 0.9690 0.02441 0.01723 0.0217 0.0122 0.7537 10.750 0.9787 0.02549 0.01836 0.0229 0.0122 0.7594 11.000 0.9883 0.02662 0.01956 0.0240 0.0122 0.7652 11.250 0.9965 0.02788 0.02088 0.0251 0.0122 0.7706 11.500 1.0056 0.02909 0.02218 0.0261 0.0121 0.7771 11.750 1.0138 0.03043 0.02358 0.0270 0.0121 0.7839 12.000 1.0223 0.03175 0.02498 0.0278 0.0120 0.7906 12.250 1.0286 0.03329 0.02661 0.0286 0.0120 0.7987 12.500 1.0366 0.03472 0.02812 0.0293 0.0119 0.8067 12.750 1.0431 0.03629 0.02980 0.0299 0.0119 0.8160 13.000 1.0498 0.03787 0.03149 0.0305 0.0119 0.8271 13.250 1.0558 0.03952 0.03331 0.0309 0.0117 0.8471 13.500 1.0899 0.04184 0.03593 0.0250 0.0116 0.9173 13.750 1.0981 0.04409 0.03826 0.0242 0.0117 1.0000 14.000 1.1034 0.04616 0.04040 0.0240 0.0116 1.0000 14.250 1.1067 0.04849 0.04280 0.0239 0.0116 1.0000 14.500 1.1106 0.05082 0.04521 0.0236 0.0116 1.0000 14.750 1.1145 0.05327 0.04773 0.0232 0.0116 1.0000 15.000 1.1164 0.05602 0.05057 0.0227 0.0115 1.0000 15.250 1.1186 0.05883 0.05346 0.0219 0.0115 1.0000 15.500 1.1193 0.06190 0.05662 0.0211 0.0115 1.0000 15.750 1.1195 0.06513 0.05995 0.0200 0.0115 1.0000 16.000 1.1192 0.06852 0.06344 0.0188 0.0115 1.0000 16.250 1.1179 0.07216 0.06717 0.0174 0.0115 1.0000 16.500 1.1151 0.07610 0.07123 0.0158 0.0115 1.0000 16.750 1.1117 0.08026 0.07550 0.0139 0.0115 1.0000 17.000 1.1077 0.08465 0.07999 0.0118 0.0114 1.0000 17.250 1.1021 0.08938 0.08484 0.0095 0.0114 1.0000 17.500 1.0954 0.09445 0.09004 0.0070 0.0115 1.0000 17.750 1.0875 0.09989 0.09560 0.0041 0.0115 1.0000 18.000 1.0783 0.10567 0.10151 0.0009 0.0114 1.0000 18.250 1.0684 0.11179 0.10776 -0.0026 0.0115 1.0000 18.500 1.0561 0.11854 0.11467 -0.0065 0.0115 1.0000 18.750 1.0427 0.12573 0.12200 -0.0108 0.0115 1.0000