XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC08-64C AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.5736 0.08143 0.07976 -0.0156 1.0000 0.0063 -8.000 -0.5812 0.07585 0.07419 -0.0216 1.0000 0.0063 -7.750 -0.5830 0.07060 0.06890 -0.0253 1.0000 0.0063 -7.500 -0.5822 0.06552 0.06375 -0.0279 1.0000 0.0063 -7.250 -0.5801 0.06073 0.05888 -0.0292 1.0000 0.0064 -7.000 -0.5779 0.05631 0.05436 -0.0290 1.0000 0.0065 -6.750 -0.5759 0.05218 0.05013 -0.0277 1.0000 0.0066 -6.500 -0.5741 0.04821 0.04603 -0.0254 1.0000 0.0068 -6.250 -0.5713 0.04316 0.04078 -0.0219 1.0000 0.0070 -6.000 -0.5605 0.03732 0.03460 -0.0197 0.9993 0.0071 -5.750 -0.5366 0.03262 0.02946 -0.0205 0.9971 0.0072 -5.500 -0.5156 0.02682 0.02319 -0.0211 0.9952 0.0073 -5.250 -0.4901 0.02451 0.02077 -0.0220 0.9928 0.0075 -5.000 -0.4622 0.02260 0.01871 -0.0229 0.9907 0.0077 -4.750 -0.4326 0.02064 0.01652 -0.0238 0.9889 0.0082 -4.500 -0.4017 0.01793 0.01326 -0.0240 0.9874 0.0098 -4.250 -0.3698 0.01660 0.01187 -0.0254 0.9864 0.0105 -4.000 -0.3402 0.01560 0.01041 -0.0253 0.9830 0.0121 -3.750 -0.3090 0.01416 0.00904 -0.0264 0.9810 0.0127 -3.500 -0.2768 0.01351 0.00809 -0.0272 0.9787 0.0148 -3.250 -0.2450 0.01238 0.00706 -0.0283 0.9767 0.0161 -3.000 -0.2168 0.01184 0.00646 -0.0285 0.9720 0.0198 -2.000 -0.1082 0.00898 0.00351 -0.0269 0.9486 0.0215 -1.750 -0.0827 0.00856 0.00304 -0.0257 0.9399 0.0135 -1.500 -0.0561 0.00828 0.00272 -0.0251 0.9326 0.0116 -1.250 -0.0311 0.00786 0.00228 -0.0244 0.9232 0.0112 -1.000 -0.0051 0.00757 0.00197 -0.0238 0.9142 0.0113 -0.750 0.0210 0.00732 0.00169 -0.0232 0.9040 0.0121 -0.500 0.0473 0.00710 0.00145 -0.0227 0.8931 0.0133 -0.250 0.0739 0.00700 0.00131 -0.0223 0.8802 0.0144 0.000 0.0994 0.00692 0.00115 -0.0215 0.8551 0.0167 0.250 0.1231 0.00651 0.00102 -0.0206 0.8266 0.1472 0.500 0.1331 0.00452 0.00086 -0.0177 0.7962 0.7491 0.750 0.1482 0.00428 0.00097 -0.0142 0.7398 0.9077 1.000 0.1669 0.00471 0.00114 -0.0115 0.6461 0.9580 1.250 0.1866 0.00591 0.00139 -0.0098 0.3973 0.9761 1.500 0.2229 0.00782 0.00191 -0.0126 0.0171 0.9826 1.750 0.2565 0.00802 0.00208 -0.0139 0.0143 0.9865 2.000 0.2914 0.00820 0.00225 -0.0155 0.0129 0.9875 2.250 0.3253 0.00847 0.00252 -0.0169 0.0120 0.9887 2.500 0.3588 0.00874 0.00278 -0.0182 0.0109 0.9902 2.750 0.3912 0.00910 0.00313 -0.0192 0.0104 0.9920 3.000 0.4227 0.00950 0.00353 -0.0201 0.0103 0.9939 3.250 0.4542 0.00996 0.00402 -0.0209 0.0108 0.9955 3.500 0.4856 0.01068 0.00474 -0.0218 0.0118 0.9966 5.500 0.6862 0.01658 0.01100 -0.0170 0.0138 1.0000 5.750 0.7066 0.01741 0.01189 -0.0156 0.0130 1.0000 6.000 0.7280 0.01880 0.01320 -0.0147 0.0125 1.0000 6.250 0.7447 0.01980 0.01462 -0.0122 0.0111 1.0000 6.500 0.7637 0.02087 0.01581 -0.0106 0.0104 1.0000 6.750 0.7826 0.02221 0.01718 -0.0092 0.0100 1.0000 7.000 0.7946 0.02638 0.02157 -0.0069 0.0095 1.0000 7.250 0.8107 0.02652 0.02224 -0.0038 0.0083 1.0000 7.500 0.8254 0.02841 0.02434 -0.0018 0.0080 1.0000 7.750 0.8395 0.03032 0.02642 0.0000 0.0078 1.0000 8.000 0.8531 0.03230 0.02856 0.0016 0.0077 1.0000 8.250 0.8653 0.03448 0.03089 0.0033 0.0075 1.0000 8.500 0.8764 0.03692 0.03343 0.0048 0.0074 1.0000 8.750 0.8678 0.04367 0.04045 0.0076 0.0072 1.0000 9.000 0.8675 0.04697 0.04403 0.0099 0.0072 1.0000 9.250 0.8632 0.05045 0.04776 0.0122 0.0072 1.0000 9.500 0.8544 0.05406 0.05159 0.0143 0.0072 1.0000 9.750 0.8387 0.05736 0.05505 0.0169 0.0072 1.0000 10.000 0.8211 0.06104 0.05888 0.0179 0.0072 1.0000 10.250 0.8025 0.06581 0.06378 0.0162 0.0072 1.0000 10.500 0.7847 0.07209 0.07019 0.0115 0.0072 1.0000 10.750 0.7675 0.08154 0.07976 0.0028 0.0072 1.0000