XFOIL Version 6.94 Calculated polar for: FX 60-177 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3193 0.01587 0.01082 -0.0773 0.8181 0.7325 1.000 0.4626 0.01477 0.00965 -0.0836 0.7964 0.7466 1.500 0.5082 0.01460 0.00949 -0.0809 0.7791 0.7511 2.000 0.5655 0.01435 0.00923 -0.0808 0.7604 0.7574 2.500 0.6400 0.01394 0.00872 -0.0857 0.7392 0.7656 3.000 0.6764 0.01381 0.00858 -0.0814 0.7108 0.7695 3.500 0.7217 0.01383 0.00851 -0.0790 0.6775 0.7739 4.000 0.7613 0.01399 0.00856 -0.0762 0.6354 0.7798 4.500 0.8126 0.01431 0.00865 -0.0765 0.5869 0.7872 5.000 0.8363 0.01482 0.00900 -0.0704 0.5402 0.7915 5.500 0.8653 0.01548 0.00951 -0.0659 0.4943 0.7962 6.000 0.8997 0.01624 0.01009 -0.0631 0.4475 0.8016 6.500 0.9371 0.01727 0.01085 -0.0616 0.3951 0.8071 7.000 0.9512 0.01844 0.01172 -0.0551 0.3376 0.8116 7.500 0.9833 0.01937 0.01259 -0.0521 0.3037 0.8163 8.000 1.0178 0.02053 0.01361 -0.0502 0.2696 0.8212 8.500 1.0539 0.02202 0.01485 -0.0492 0.2247 0.8261 9.000 1.0815 0.02319 0.01596 -0.0459 0.1970 0.8304 9.500 1.1118 0.02452 0.01727 -0.0435 0.1759 0.8350 10.000 1.1468 0.02592 0.01868 -0.0421 0.1564 0.8402 10.500 1.1775 0.02801 0.02057 -0.0411 0.1244 0.8448 11.000 1.2027 0.02975 0.02232 -0.0385 0.1036 0.8485 11.500 1.2270 0.03183 0.02439 -0.0363 0.0867 0.8527 12.000 1.2522 0.03421 0.02680 -0.0349 0.0709 0.8572 12.500 1.2724 0.03739 0.02991 -0.0337 0.0533 0.8614 13.500 1.2962 0.04436 0.03699 -0.0296 0.0289 0.8704 14.000 1.3118 0.04822 0.04101 -0.0290 0.0230 0.8748 14.500 1.2967 0.05591 0.04870 -0.0286 0.0027 0.8783 15.000 1.2976 0.06171 0.05474 -0.0285 0.0018 0.8819 15.500 1.2990 0.06791 0.06120 -0.0292 0.0016 0.8859 16.000 1.2986 0.07508 0.06866 -0.0312 0.0015 0.8898 16.500 1.2953 0.08334 0.07721 -0.0343 0.0015 0.8934 17.000 1.2861 0.09236 0.08654 -0.0376 0.0014 0.8969 17.500 1.2728 0.10214 0.09663 -0.0416 0.0014 0.9007 18.000 1.2585 0.11286 0.10765 -0.0468 0.0014 0.9043 18.500 1.2435 0.12429 0.11939 -0.0532 0.0013 0.9073 19.000 1.2279 0.13608 0.13147 -0.0602 0.0013 0.9103 19.500 1.2097 0.14764 0.14332 -0.0668 0.0013 0.9143 20.000 1.1927 0.15952 0.15546 -0.0742 0.0014 0.9190 20.500 1.1774 0.17184 0.16804 -0.0826 0.0014 0.9234 21.000 1.1599 0.18379 0.18025 -0.0903 0.0014 0.9285 21.500 1.1432 0.19624 0.19292 -0.0986 0.0014 0.9348 22.000 1.1253 0.20849 0.20539 -0.1064 0.0014 0.9430 22.500 1.1070 0.21994 0.21704 -0.1130 0.0015 0.9581 23.000 1.0986 0.23078 0.22804 -0.1210 0.0015 1.0000 23.500 1.1011 0.24205 0.23943 -0.1304 0.0015 1.0000 24.000 1.1090 0.25174 0.24921 -0.1389 0.0016 1.0000 24.500 1.1201 0.26011 0.25766 -0.1466 0.0016 1.0000 25.000 1.1329 0.26747 0.26507 -0.1537 0.0016 1.0000 25.500 1.1464 0.27412 0.27177 -0.1602 0.0016 1.0000 26.000 1.1593 0.28053 0.27824 -0.1665 0.0016 1.0000 26.500 1.1710 0.28705 0.28483 -0.1727 0.0016 1.0000 27.000 1.1792 0.29456 0.29243 -0.1792 0.0017 1.0000 27.500 1.1657 0.31355 0.31158 -0.1891 0.0020 1.0000 28.000 0.8271 0.36360 0.36221 -0.1578 0.0103 0.9012 28.500 0.8331 0.37193 0.37057 -0.1609 0.0097 0.9051 29.000 0.8420 0.38022 0.37889 -0.1646 0.0094 0.9088 29.500 0.8453 0.38924 0.38792 -0.1674 0.0093 0.9119 30.000 0.8465 0.39957 0.39826 -0.1703 0.0088 0.9151 30.500 0.8506 0.40848 0.40720 -0.1734 0.0081 0.9188 31.000 0.8564 0.41805 0.41680 -0.1771 0.0077 0.9224 31.500 0.8596 0.42524 0.42402 -0.1797 0.0073 0.9264 32.000 0.8615 0.43058 0.42938 -0.1813 0.0071 0.9316 32.500 0.8646 0.44066 0.43947 -0.1847 0.0070 0.9359 33.000 0.8645 0.44890 0.44772 -0.1869 0.0069 0.9412 33.500 0.8645 0.45608 0.45492 -0.1888 0.0064 0.9483