XFOIL Version 6.96 Calculated polar for: FX 66-S-161 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 ------ -------- --------- --------- -------- -------- -------- -11.250 -0.3764 0.06081 0.05841 -0.0908 0.9105 0.0157 -11.000 -0.3965 0.04925 0.04620 -0.1068 0.8359 0.0155 -10.750 -0.4353 0.04294 0.03945 -0.1093 0.7925 0.0147 -10.500 -0.5235 0.03369 0.02928 -0.1067 0.7656 0.0134 -10.250 -0.5258 0.03158 0.02686 -0.1048 0.7469 0.0134 -10.000 -0.5210 0.02975 0.02478 -0.1033 0.7317 0.0136 -9.750 -0.5126 0.02791 0.02269 -0.1020 0.7188 0.0138 -9.500 -0.5005 0.02623 0.02077 -0.1009 0.7075 0.0139 -9.250 -0.4855 0.02481 0.01913 -0.0999 0.6976 0.0142 -9.000 -0.4674 0.02376 0.01792 -0.0991 0.6878 0.0145 -8.750 -0.4471 0.02306 0.01709 -0.0985 0.6789 0.0153 -8.500 -0.4276 0.02188 0.01570 -0.0977 0.6708 0.0157 -8.250 -0.4070 0.02069 0.01431 -0.0970 0.6630 0.0161 -8.000 -0.3857 0.01965 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1.1796 0.01145 0.00619 -0.0965 0.4731 1.0000 6.750 1.2057 0.01155 0.00637 -0.0964 0.4698 1.0000 7.000 1.2323 0.01167 0.00653 -0.0963 0.4664 1.0000 7.250 1.2590 0.01178 0.00666 -0.0963 0.4629 1.0000 7.500 1.2840 0.01187 0.00682 -0.0959 0.4579 1.0000 7.750 1.3076 0.01180 0.00680 -0.0952 0.4500 1.0000 8.000 1.3295 0.01172 0.00677 -0.0942 0.4379 1.0000 8.250 1.3513 0.01173 0.00680 -0.0932 0.4235 1.0000 8.500 1.3735 0.01184 0.00694 -0.0924 0.4097 1.0000 8.750 1.3953 0.01204 0.00716 -0.0915 0.3968 1.0000 9.000 1.4145 0.01236 0.00744 -0.0903 0.3796 1.0000 9.250 1.4301 0.01283 0.00785 -0.0884 0.3567 1.0000 9.500 1.4396 0.01356 0.00845 -0.0857 0.3253 1.0000 9.750 1.4406 0.01451 0.00923 -0.0815 0.2913 1.0000 10.000 1.4279 0.01592 0.01039 -0.0754 0.2472 1.0000 10.250 1.4108 0.01770 0.01195 -0.0694 0.2057 1.0000 10.500 1.3905 0.02000 0.01406 -0.0641 0.1691 1.0000 10.750 1.3713 0.02278 0.01670 -0.0602 0.1392 1.0000 11.000 1.3510 0.02618 0.01997 -0.0573 0.1099 1.0000 11.250 1.3331 0.02981 0.02349 -0.0552 0.0846 1.0000 11.500 1.3152 0.03368 0.02726 -0.0535 0.0619 1.0000 11.750 1.3004 0.03741 0.03092 -0.0522 0.0442 1.0000 12.000 1.2883 0.04101 0.03448 -0.0511 0.0324 1.0000 12.250 1.2809 0.04425 0.03773 -0.0503 0.0262 1.0000 12.500 1.2776 0.04720 0.04073 -0.0498 0.0225 1.0000 12.750 1.2724 0.05044 0.04400 -0.0493 0.0200 1.0000 13.000 1.2707 0.05344 0.04708 -0.0490 0.0181 1.0000 13.250 1.2709 0.05629 0.04999 -0.0489 0.0167 1.0000 13.500 1.2682 0.05953 0.05328 -0.0488 0.0154 1.0000 13.750 1.2618 0.06332 0.05716 -0.0489 0.0142 1.0000 14.000 1.2640 0.06614 0.06007 -0.0490 0.0131 1.0000 14.250 1.2641 0.06926 0.06327 -0.0492 0.0128 1.0000 14.500 1.2638 0.07251 0.06657 -0.0495 0.0120 1.0000 14.750 1.2616 0.07604 0.07017 -0.0498 0.0114 1.0000 15.000 1.2548 0.08026 0.07446 -0.0504 0.0109 1.0000 15.250 1.2495 0.08430 0.07858 -0.0509 0.0105 1.0000 15.500 1.2528 0.08728 0.08164 -0.0514 0.0102 1.0000