XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC-08(B)3 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.250 -0.6342 0.08502 0.08279 -0.0067 1.0000 0.0106 -9.000 -0.6442 0.07828 0.07607 -0.0135 1.0000 0.0106 -8.750 -0.6563 0.07194 0.06969 -0.0179 1.0000 0.0105 -8.500 -0.6636 0.06510 0.06275 -0.0213 1.0000 0.0105 -8.250 -0.6686 0.05816 0.05566 -0.0232 1.0000 0.0106 -8.000 -0.6715 0.05117 0.04844 -0.0235 1.0000 0.0107 -7.750 -0.6737 0.04376 0.04071 -0.0224 1.0000 0.0109 -7.500 -0.6797 0.03488 0.03129 -0.0193 1.0000 0.0111 -7.250 -0.6813 0.02825 0.02402 -0.0153 1.0000 0.0114 -7.000 -0.6714 0.02545 0.02084 -0.0125 1.0000 0.0118 -6.750 -0.6517 0.02314 0.01815 -0.0115 0.9977 0.0120 -6.500 -0.6241 0.02115 0.01581 -0.0120 0.9945 0.0122 -6.250 -0.5984 0.01886 0.01313 -0.0121 0.9902 0.0125 -6.000 -0.5708 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-0.0187 0.1827 0.9967 3.750 0.5364 0.00979 0.00296 -0.0194 0.1315 0.9978 4.000 0.5656 0.01018 0.00318 -0.0201 0.0933 0.9988 4.250 0.5945 0.01056 0.00344 -0.0207 0.0649 0.9997 4.500 0.6199 0.01087 0.00368 -0.0204 0.0497 1.0000 4.750 0.6438 0.01114 0.00392 -0.0198 0.0415 1.0000 5.000 0.6675 0.01140 0.00419 -0.0191 0.0359 1.0000 5.250 0.6908 0.01170 0.00449 -0.0183 0.0314 1.0000 5.500 0.7138 0.01205 0.00486 -0.0175 0.0279 1.0000 5.750 0.7369 0.01234 0.00520 -0.0167 0.0262 1.0000 6.000 0.7597 0.01266 0.00555 -0.0158 0.0242 1.0000 6.250 0.7819 0.01305 0.00593 -0.0148 0.0223 1.0000 6.500 0.8027 0.01363 0.00658 -0.0137 0.0206 1.0000 6.750 0.8245 0.01402 0.00703 -0.0126 0.0200 1.0000 7.000 0.8458 0.01446 0.00754 -0.0115 0.0193 1.0000 7.250 0.8669 0.01491 0.00805 -0.0103 0.0185 1.0000 7.500 0.8879 0.01534 0.00853 -0.0092 0.0175 1.0000 7.750 0.9087 0.01577 0.00902 -0.0081 0.0166 1.0000 8.000 0.9284 0.01634 0.00963 -0.0068 0.0158 1.0000 8.250 0.9454 0.01725 0.01061 -0.0050 0.0148 1.0000 8.500 0.9660 0.01767 0.01112 -0.0039 0.0143 1.0000 8.750 0.9859 0.01818 0.01171 -0.0026 0.0136 1.0000 9.000 1.0049 0.01878 0.01240 -0.0012 0.0129 1.0000 9.250 1.0237 0.01938 0.01310 0.0002 0.0123 1.0000 9.500 1.0425 0.01997 0.01376 0.0015 0.0117 1.0000 9.750 1.0605 0.02062 0.01447 0.0029 0.0112 1.0000 10.000 1.0740 0.02183 0.01578 0.0049 0.0105 1.0000 10.250 1.0921 0.02248 0.01656 0.0063 0.0101 1.0000 10.500 1.1094 0.02323 0.01745 0.0077 0.0096 1.0000 10.750 1.1261 0.02405 0.01840 0.0092 0.0091 1.0000 11.000 1.1421 0.02491 0.01940 0.0107 0.0086 1.0000 11.250 1.1575 0.02579 0.02040 0.0121 0.0082 1.0000 11.500 1.1725 0.02667 0.02138 0.0136 0.0079 1.0000 11.750 1.1863 0.02760 0.02243 0.0151 0.0075 1.0000 12.000 1.1963 0.02886 0.02381 0.0170 0.0073 1.0000 12.250 1.1972 0.03068 0.02580 0.0200 0.0070 1.0000 12.500 1.2004 0.03223 0.02755 0.0226 0.0069 1.0000 12.750 1.2011 0.03409 0.02961 0.0249 0.0067 1.0000 13.000 1.1995 0.03625 0.03198 0.0269 0.0066 1.0000 13.250 1.1953 0.03881 0.03476 0.0284 0.0065 1.0000 13.500 1.1877 0.04192 0.03809 0.0293 0.0064 1.0000 13.750 1.1778 0.04561 0.04200 0.0293 0.0063 1.0000 14.000 1.1644 0.05020 0.04681 0.0281 0.0062 1.0000 14.250 1.1470 0.05603 0.05286 0.0254 0.0062 1.0000 14.500 1.1255 0.06359 0.06063 0.0206 0.0062 1.0000 14.750 1.0974 0.07382 0.07108 0.0134 0.0063 1.0000 15.000 1.0587 0.08753 0.08499 0.0045 0.0064 1.0000 15.250 1.0025 0.10533 0.10293 -0.0052 0.0067 1.0000