XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC10-64C 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 ------ -------- --------- --------- -------- -------- -------- -10.500 -0.6411 0.08136 0.07765 -0.0310 1.0000 0.0206 -10.250 -0.6924 0.06546 0.06168 -0.0431 1.0000 0.0200 -10.000 -0.7547 0.05446 0.05036 -0.0416 1.0000 0.0196 -9.750 -0.7843 0.04832 0.04387 -0.0379 1.0000 0.0196 -9.500 -0.7974 0.04413 0.03935 -0.0343 1.0000 0.0198 -9.250 -0.8030 0.04066 0.03554 -0.0309 1.0000 0.0200 -9.000 -0.8031 0.03771 0.03226 -0.0276 1.0000 0.0204 -8.750 -0.7984 0.03539 0.02964 -0.0247 1.0000 0.0210 -8.500 -0.7921 0.03309 0.02700 -0.0218 1.0000 0.0217 -8.250 -0.7840 0.03084 0.02436 -0.0189 1.0000 0.0224 -8.000 -0.7732 0.02879 0.02193 -0.0162 1.0000 0.0228 -7.750 -0.7597 0.02704 0.01983 -0.0139 1.0000 0.0232 -7.500 -0.7441 0.02561 0.01810 -0.0119 1.0000 0.0236 -7.250 -0.7280 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0.00381 -0.0255 0.6838 0.9669 2.750 0.4526 0.01009 0.00386 -0.0253 0.6457 0.9721 3.000 0.4821 0.01040 0.00387 -0.0255 0.5769 0.9754 3.250 0.5077 0.01089 0.00396 -0.0252 0.4881 0.9798 3.500 0.5325 0.01147 0.00415 -0.0248 0.4016 0.9844 3.750 0.5589 0.01223 0.00445 -0.0252 0.3005 0.9880 4.000 0.5849 0.01302 0.00481 -0.0255 0.2137 0.9921 4.250 0.6139 0.01365 0.00517 -0.0263 0.1558 0.9955 4.500 0.6442 0.01423 0.00557 -0.0274 0.1171 0.9989 4.750 0.6666 0.01473 0.00594 -0.0267 0.0914 1.0000 5.000 0.6850 0.01517 0.00632 -0.0250 0.0769 1.0000 5.250 0.7034 0.01561 0.00673 -0.0234 0.0675 1.0000 5.500 0.7220 0.01602 0.00717 -0.0217 0.0619 1.0000 5.750 0.7402 0.01648 0.00764 -0.0200 0.0570 1.0000 6.000 0.7572 0.01703 0.00818 -0.0181 0.0530 1.0000 6.250 0.7755 0.01747 0.00869 -0.0163 0.0500 1.0000 6.500 0.7932 0.01798 0.00924 -0.0145 0.0476 1.0000 6.750 0.8105 0.01855 0.00984 -0.0127 0.0455 1.0000 7.000 0.8263 0.01929 0.01060 -0.0107 0.0434 1.0000 7.250 0.8429 0.02003 0.01138 -0.0088 0.0418 1.0000 7.500 0.8610 0.02061 0.01206 -0.0071 0.0401 1.0000 7.750 0.8789 0.02129 0.01281 -0.0054 0.0384 1.0000 8.000 0.8968 0.02204 0.01363 -0.0038 0.0371 1.0000 8.250 0.9147 0.02283 0.01449 -0.0023 0.0359 1.0000 8.500 0.9325 0.02370 0.01541 -0.0008 0.0348 1.0000 8.750 0.9497 0.02493 0.01667 0.0006 0.0335 1.0000 9.000 0.9681 0.02603 0.01790 0.0019 0.0325 1.0000 9.250 0.9866 0.02694 0.01898 0.0033 0.0313 1.0000 9.500 1.0044 0.02806 0.02030 0.0046 0.0303 1.0000 9.750 1.0214 0.02929 0.02170 0.0061 0.0293 1.0000 10.000 1.0375 0.03054 0.02311 0.0075 0.0284 1.0000 10.250 1.0525 0.03165 0.02435 0.0090 0.0275 1.0000 10.500 1.0666 0.03268 0.02546 0.0106 0.0267 1.0000 10.750 1.0787 0.03412 0.02697 0.0122 0.0258 1.0000 11.000 1.0875 0.03596 0.02907 0.0143 0.0251 1.0000 11.250 1.0935 0.03778 0.03122 0.0168 0.0245 1.0000 11.500 1.0949 0.03971 0.03344 0.0198 0.0238 1.0000 11.750 1.0938 0.04172 0.03571 0.0227 0.0231 1.0000 12.000 1.0913 0.04377 0.03799 0.0252 0.0225 1.0000 12.250 1.0884 0.04581 0.04024 0.0274 0.0219 1.0000 12.500 1.0845 0.04800 0.04261 0.0292 0.0214 1.0000 12.750 1.0812 0.05020 0.04496 0.0304 0.0209 1.0000 13.000 1.0758 0.05284 0.04774 0.0311 0.0206 1.0000 13.250 1.0690 0.05584 0.05089 0.0313 0.0203 1.0000 13.500 1.0583 0.05963 0.05484 0.0308 0.0201 1.0000 13.750 1.0436 0.06434 0.05973 0.0293 0.0200 1.0000 14.000 1.0235 0.07042 0.06601 0.0264 0.0199 1.0000 14.250 0.9808 0.08176 0.07770 0.0192 0.0201 1.0000