XFOIL Version 6.96 Calculated polar for: RAE 101 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.8410 0.05365 0.04905 -0.0242 1.0000 0.0364 -9.250 -0.8464 0.04864 0.04383 -0.0228 1.0000 0.0354 -9.000 -0.8474 0.04473 0.03957 -0.0209 1.0000 0.0354 -8.750 -0.8444 0.04081 0.03525 -0.0189 1.0000 0.0351 -8.500 -0.8381 0.03685 0.03088 -0.0169 1.0000 0.0347 -8.250 -0.8274 0.03330 0.02691 -0.0149 1.0000 0.0343 -8.000 -0.8128 0.03034 0.02355 -0.0131 1.0000 0.0345 -7.750 -0.7950 0.02796 0.02080 -0.0115 1.0000 0.0353 -7.500 -0.7748 0.02656 0.01908 -0.0102 1.0000 0.0367 -7.250 -0.7555 0.02376 0.01607 -0.0090 1.0000 0.0388 -7.000 -0.7341 0.02221 0.01448 -0.0080 1.0000 0.0409 -6.750 -0.7121 0.02094 0.01309 -0.0069 1.0000 0.0433 -6.500 -0.6900 0.01989 0.01192 -0.0058 1.0000 0.0464 -6.250 -0.6701 0.01842 0.01043 -0.0044 1.0000 0.0504 -6.000 -0.6498 0.01748 0.00950 -0.0031 1.0000 0.0551 -5.750 -0.6293 0.01663 0.00857 -0.0016 1.0000 0.0604 -5.500 -0.6109 0.01564 0.00763 0.0000 1.0000 0.0694 -5.250 -0.5924 0.01470 0.00674 0.0016 1.0000 0.0805 -5.000 -0.5737 0.01383 0.00595 0.0032 1.0000 0.1014 -4.750 -0.5572 0.01269 0.00516 0.0050 1.0000 0.1546 -4.500 -0.5448 0.01124 0.00455 0.0071 1.0000 0.3009 -4.250 -0.5275 0.01062 0.00434 0.0089 1.0000 0.4089 -4.000 -0.5077 0.01032 0.00420 0.0105 1.0000 0.4644 -3.750 -0.4871 0.01011 0.00409 0.0119 1.0000 0.5047 -3.500 -0.4665 0.00995 0.00400 0.0134 1.0000 0.5377 -3.250 -0.4460 0.00982 0.00392 0.0149 1.0000 0.5672 -3.000 -0.4258 0.00972 0.00387 0.0163 1.0000 0.5940 -2.750 -0.4060 0.00964 0.00385 0.0179 1.0000 0.6200 -2.500 -0.3867 0.00960 0.00387 0.0194 1.0000 0.6456 -2.250 -0.3637 0.00957 0.00392 0.0202 0.9985 0.6717 -2.000 -0.3200 0.00953 0.00396 0.0169 0.9890 0.7030 -1.750 -0.2773 0.00945 0.00398 0.0139 0.9796 0.7324 -1.500 -0.2339 0.00936 0.00396 0.0109 0.9708 0.7590 -1.250 -0.1934 0.00923 0.00391 0.0085 0.9600 0.7836 -1.000 -0.1514 0.00911 0.00385 0.0059 0.9502 0.8078 -0.750 -0.1092 0.00898 0.00380 0.0034 0.9402 0.8290 -0.500 -0.0712 0.00888 0.00374 0.0018 0.9266 0.8516 -0.250 -0.0347 0.00882 0.00372 0.0007 0.9105 0.8717 0.000 0.0000 0.00879 0.00371 0.0000 0.8914 0.8914 0.250 0.0347 0.00882 0.00372 -0.0007 0.8717 0.9105 0.500 0.0711 0.00888 0.00374 -0.0018 0.8515 0.9265 0.750 0.1092 0.00898 0.00380 -0.0034 0.8291 0.9402 1.000 0.1513 0.00911 0.00385 -0.0059 0.8078 0.9502 1.250 0.1934 0.00923 0.00391 -0.0085 0.7835 0.9600 1.500 0.2339 0.00935 0.00396 -0.0109 0.7592 0.9708 1.750 0.2774 0.00945 0.00398 -0.0139 0.7325 0.9796 2.000 0.3200 0.00952 0.00396 -0.0169 0.7032 0.9891 2.250 0.3636 0.00956 0.00392 -0.0202 0.6717 0.9986 2.500 0.3865 0.00960 0.00387 -0.0194 0.6457 1.0000 2.750 0.4058 0.00964 0.00385 -0.0178 0.6199 1.0000 3.000 0.4256 0.00971 0.00387 -0.0163 0.5941 1.0000 3.250 0.4458 0.00982 0.00392 -0.0148 0.5674 1.0000 3.500 0.4663 0.00995 0.00400 -0.0134 0.5380 1.0000 3.750 0.4869 0.01011 0.00409 -0.0119 0.5045 1.0000 4.000 0.5075 0.01031 0.00420 -0.0104 0.4649 1.0000 4.250 0.5273 0.01062 0.00434 -0.0089 0.4090 1.0000 4.500 0.5446 0.01123 0.00455 -0.0071 0.3012 1.0000 4.750 0.5571 0.01269 0.00516 -0.0050 0.1543 1.0000 5.000 0.5736 0.01383 0.00595 -0.0032 0.1015 1.0000 5.250 0.5923 0.01471 0.00674 -0.0016 0.0807 1.0000 5.500 0.6108 0.01564 0.00763 0.0001 0.0693 1.0000 5.750 0.6292 0.01663 0.00856 0.0016 0.0604 1.0000 6.000 0.6498 0.01748 0.00950 0.0031 0.0551 1.0000 6.250 0.6701 0.01843 0.01043 0.0044 0.0504 1.0000 6.500 0.6900 0.01989 0.01192 0.0058 0.0464 1.0000 6.750 0.7121 0.02094 0.01309 0.0069 0.0433 1.0000 7.000 0.7341 0.02223 0.01449 0.0080 0.0410 1.0000 7.250 0.7556 0.02375 0.01606 0.0090 0.0389 1.0000 7.500 0.7748 0.02665 0.01917 0.0101 0.0367 1.0000 7.750 0.7951 0.02796 0.02080 0.0115 0.0353 1.0000 8.000 0.8129 0.03035 0.02355 0.0131 0.0346 1.0000 8.250 0.8276 0.03328 0.02688 0.0149 0.0344 1.0000 8.500 0.8383 0.03685 0.03088 0.0168 0.0347 1.0000 8.750 0.8447 0.04080 0.03524 0.0188 0.0352 1.0000 9.000 0.8477 0.04475 0.03958 0.0208 0.0355 1.0000 9.250 0.8467 0.04895 0.04410 0.0226 0.0358 1.0000 9.500 0.8416 0.05332 0.04874 0.0242 0.0362 1.0000 12.250 0.5664 0.12453 0.12116 -0.0012 0.0597 1.0000 12.500 0.5552 0.12989 0.12649 -0.0053 0.0594 1.0000