XFOIL Version 6.96 Calculated polar for: WORTMANN FX 66-17A-175 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 ------ -------- --------- --------- -------- -------- -------- -11.750 -0.2238 0.09268 0.08792 -0.0794 0.6516 0.0124 -11.500 -0.2311 0.08630 0.08154 -0.0824 0.6467 0.0121 -11.250 -0.2568 0.07452 0.06975 -0.0886 0.6456 0.0117 -11.000 -0.2843 0.06557 0.06068 -0.0946 0.6429 0.0115 -10.750 -0.3075 0.05908 0.05405 -0.0989 0.6392 0.0114 -10.500 -0.3256 0.05406 0.04888 -0.1020 0.6345 0.0111 -10.250 -0.3442 0.04944 0.04407 -0.1044 0.6300 0.0111 -10.000 -0.3604 0.04546 0.03987 -0.1061 0.6257 0.0111 -9.750 -0.3729 0.04224 0.03646 -0.1068 0.6208 0.0110 -9.500 -0.3863 0.03965 0.03362 -0.1056 0.6162 0.0111 -9.250 -0.3892 0.03707 0.03072 -0.1046 0.6119 0.0111 -9.000 -0.3862 0.03452 0.02785 -0.1036 0.6069 0.0113 -8.750 -0.3780 0.03223 0.02522 -0.1026 0.6019 0.0115 -8.500 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0.01564 0.00616 -0.0933 0.4860 0.5451 1.250 0.6034 0.01573 0.00619 -0.0935 0.4844 0.5474 1.500 0.6313 0.01582 0.00625 -0.0936 0.4829 0.5491 1.750 0.6585 0.01591 0.00638 -0.0937 0.4810 0.5509 2.000 0.6852 0.01601 0.00652 -0.0937 0.4786 0.5529 2.250 0.7121 0.01610 0.00665 -0.0937 0.4762 0.5552 2.500 0.7391 0.01620 0.00676 -0.0938 0.4739 0.5579 2.750 0.7664 0.01631 0.00688 -0.0939 0.4720 0.5607 3.000 0.7937 0.01642 0.00699 -0.0940 0.4704 0.5632 3.250 0.8211 0.01654 0.00713 -0.0941 0.4690 0.5656 3.500 0.8482 0.01669 0.00733 -0.0942 0.4675 0.5684 3.750 0.8741 0.01687 0.00760 -0.0941 0.4656 0.5713 4.000 0.9001 0.01704 0.00785 -0.0940 0.4636 0.5743 4.250 0.9263 0.01721 0.00809 -0.0940 0.4616 0.5772 4.500 0.9525 0.01735 0.00832 -0.0940 0.4597 0.5800 4.750 0.9789 0.01751 0.00854 -0.0939 0.4581 0.5834 5.000 1.0056 0.01768 0.00877 -0.0940 0.4566 0.5873 5.250 1.0325 0.01787 0.00901 -0.0941 0.4554 0.5913 5.500 1.0595 0.01805 0.00926 -0.0942 0.4542 0.5948 5.750 1.0840 0.01831 0.00968 -0.0939 0.4525 0.5989 6.000 1.1082 0.01860 0.01011 -0.0937 0.4507 0.6036 6.250 1.1327 0.01887 0.01055 -0.0934 0.4489 0.6084 6.500 1.1572 0.01914 0.01096 -0.0932 0.4472 0.6138 6.750 1.1820 0.01941 0.01135 -0.0929 0.4456 0.6203 7.000 1.2068 0.01965 0.01174 -0.0927 0.4441 0.6271 7.250 1.2321 0.01988 0.01211 -0.0926 0.4427 0.6359 7.500 1.2579 0.02008 0.01244 -0.0925 0.4412 0.6465 7.750 1.2847 0.02031 0.01280 -0.0926 0.4400 0.6599 8.000 1.3055 0.02070 0.01345 -0.0918 0.4380 0.6770 8.250 1.3227 0.02063 0.01364 -0.0901 0.4310 0.7021 8.500 1.3381 0.01979 0.01300 -0.0876 0.4167 0.7564 8.750 1.3581 0.01923 0.01281 -0.0863 0.4017 1.0000 9.000 1.3695 0.01916 0.01269 -0.0836 0.3830 1.0000 9.250 1.3704 0.01954 0.01306 -0.0793 0.3517 1.0000 9.750 1.3250 0.02307 0.01597 -0.0665 0.2528 1.0000 10.000 1.2965 0.02636 0.01912 -0.0621 0.2190 1.0000 10.250 1.2681 0.03064 0.02332 -0.0596 0.1932 1.0000 10.500 1.2467 0.03495 0.02760 -0.0583 0.1710 1.0000 10.750 1.2173 0.04023 0.03279 -0.0572 0.1457 1.0000 11.000 1.1969 0.04476 0.03727 -0.0563 0.1249 1.0000 11.250 1.1696 0.05006 0.04244 -0.0556 0.0973 1.0000 11.500 1.1427 0.05548 0.04769 -0.0550 0.0652 1.0000 11.750 1.1167 0.06112 0.05311 -0.0548 0.0301 1.0000 12.000 1.1108 0.06475 0.05671 -0.0548 0.0199 1.0000 12.250 1.1056 0.06839 0.06036 -0.0548 0.0136 1.0000 12.500 1.1037 0.07172 0.06374 -0.0549 0.0115 1.0000 12.750 1.1056 0.07466 0.06679 -0.0550 0.0103 1.0000 13.000 1.1080 0.07758 0.06981 -0.0552 0.0097 1.0000 13.250 1.1098 0.08058 0.07290 -0.0554 0.0091 1.0000 13.500 1.1107 0.08376 0.07618 -0.0557 0.0086 1.0000 13.750 1.1108 0.08709 0.07960 -0.0561 0.0082 1.0000 14.000 1.1099 0.09060 0.08321 -0.0566 0.0080 1.0000 14.250 1.1089 0.09417 0.08689 -0.0571 0.0077 1.0000