XFOIL Version 6.96 Calculated polar for: Eh 1.0/7.0 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.6028 0.09476 0.09022 0.0068 1.0000 0.0848 -8.000 -0.6175 0.09086 0.08640 -0.0003 1.0000 0.0860 -7.750 -0.6320 0.08730 0.08270 -0.0078 1.0000 0.0868 -7.500 -0.6064 0.08181 0.07740 -0.0008 1.0000 0.0909 -7.250 -0.6023 0.07787 0.07345 -0.0027 1.0000 0.0953 -7.000 -0.6167 0.07505 0.07014 -0.0119 1.0000 0.1002 -6.750 -0.5963 0.06877 0.06424 -0.0083 1.0000 0.1038 -6.500 -0.5963 0.06636 0.06130 -0.0129 1.0000 0.1138 -6.250 -0.5786 0.06073 0.05602 -0.0107 1.0000 0.1176 -6.000 -0.5690 0.05694 0.05199 -0.0123 1.0000 0.1290 -5.750 -0.5562 0.05357 0.04844 -0.0127 1.0000 0.1419 -5.500 -0.5415 0.05009 0.04500 -0.0119 1.0000 0.1571 -5.250 -0.5280 0.04730 0.04207 -0.0113 1.0000 0.1833 -4.750 -0.4668 0.03319 0.02599 -0.0116 1.0000 0.0709 -4.500 -0.4413 0.02902 0.02101 -0.0099 1.0000 0.0586 -4.250 -0.4169 0.02573 0.01731 -0.0087 1.0000 0.0556 -4.000 -0.3909 0.02306 0.01420 -0.0075 1.0000 0.0542 -3.750 -0.3644 0.02092 0.01175 -0.0064 1.0000 0.0552 -3.500 -0.3381 0.01938 0.00994 -0.0053 1.0000 0.0612 -3.250 -0.3140 0.01770 0.00833 -0.0043 1.0000 0.0713 -3.000 -0.2913 0.01609 0.00679 -0.0027 1.0000 0.0863 -2.750 -0.2734 0.01396 0.00536 -0.0008 1.0000 0.1746 -2.500 -0.1819 0.01164 0.00560 -0.0072 1.0000 0.9750 -2.250 -0.1089 0.01136 0.00479 -0.0164 1.0000 1.0000 -2.000 -0.0894 0.01114 0.00441 -0.0155 1.0000 1.0000 -1.750 -0.0699 0.01097 0.00406 -0.0144 1.0000 1.0000 -1.500 -0.0506 0.01082 0.00380 -0.0133 1.0000 1.0000 -1.250 -0.0319 0.01071 0.00359 -0.0121 1.0000 1.0000 -1.000 -0.0140 0.01063 0.00345 -0.0107 1.0000 1.0000 -0.750 0.0024 0.01058 0.00336 -0.0090 1.0000 1.0000 -0.500 0.0165 0.01058 0.00333 -0.0069 1.0000 1.0000 -0.250 0.0271 0.01065 0.00336 -0.0044 1.0000 1.0000 0.000 0.0326 0.01084 0.00353 -0.0013 1.0000 1.0000 0.250 0.0343 0.01117 0.00383 0.0021 1.0000 1.0000 0.500 0.1026 0.01137 0.00403 -0.0067 0.9804 1.0000 0.750 0.1699 0.01140 0.00412 -0.0150 0.9594 1.0000 1.000 0.2292 0.01134 0.00411 -0.0211 0.9325 1.0000 1.250 0.2748 0.01130 0.00409 -0.0241 0.9007 1.0000 1.500 0.3025 0.01138 0.00414 -0.0234 0.8645 1.0000 1.750 0.3251 0.01152 0.00427 -0.0216 0.8317 1.0000 2.000 0.3455 0.01171 0.00441 -0.0194 0.7993 1.0000 2.250 0.3660 0.01192 0.00456 -0.0174 0.7693 1.0000 2.500 0.3870 0.01216 0.00474 -0.0154 0.7406 1.0000 2.750 0.4084 0.01241 0.00494 -0.0136 0.7125 1.0000 3.000 0.4304 0.01268 0.00524 -0.0120 0.6850 1.0000 3.250 0.4529 0.01297 0.00551 -0.0106 0.6579 1.0000 3.500 0.4758 0.01326 0.00581 -0.0092 0.6308 1.0000 3.750 0.4987 0.01357 0.00612 -0.0079 0.6031 1.0000 4.000 0.5218 0.01387 0.00645 -0.0065 0.5748 1.0000 4.250 0.5450 0.01417 0.00682 -0.0053 0.5442 1.0000 4.500 0.5679 0.01444 0.00722 -0.0039 0.5107 1.0000 4.750 0.5866 0.01433 0.00695 -0.0015 0.4406 1.0000 5.000 0.6060 0.01461 0.00701 0.0003 0.3353 1.0000 5.250 0.6149 0.01824 0.00883 0.0026 0.0724 1.0000 5.500 0.6346 0.01995 0.01059 0.0042 0.0587 1.0000 5.750 0.6544 0.02169 0.01230 0.0056 0.0496 1.0000 6.000 0.6769 0.02342 0.01408 0.0069 0.0452 1.0000 6.250 0.7005 0.02565 0.01637 0.0080 0.0432 1.0000 6.500 0.7251 0.02827 0.01927 0.0091 0.0426 1.0000 6.750 0.7494 0.03113 0.02254 0.0105 0.0434 1.0000 7.000 0.7703 0.03476 0.02689 0.0121 0.0460 1.0000 7.250 0.7876 0.03857 0.03127 0.0135 0.0473 1.0000 7.500 0.8015 0.04243 0.03565 0.0146 0.0473 1.0000 7.750 0.8122 0.04705 0.04071 0.0155 0.0496 1.0000 8.000 0.8155 0.05353 0.04794 0.0164 0.0609 1.0000 10.000 0.7425 0.11089 0.10624 -0.0104 0.1296 1.0000 10.250 0.7083 0.11567 0.11079 -0.0196 0.1254 1.0000 10.500 0.6386 0.11525 0.11079 -0.0094 0.1298 1.0000