XFOIL Version 6.96 Calculated polar for: NASA SC(2)-0403 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 ------ -------- --------- --------- -------- -------- -------- -8.750 -0.6643 0.10961 0.10607 0.0218 1.0000 0.0135 -8.500 -0.6600 0.10560 0.10209 0.0196 1.0000 0.0135 -8.250 -0.6557 0.10156 0.09807 0.0172 1.0000 0.0135 -8.000 -0.6513 0.09744 0.09398 0.0143 1.0000 0.0136 -7.750 -0.6438 0.09286 0.08942 0.0099 1.0000 0.0136 -7.500 -0.6330 0.08776 0.08431 0.0041 1.0000 0.0136 -7.250 -0.6206 0.08261 0.07912 -0.0012 1.0000 0.0136 -7.000 -0.6071 0.07755 0.07401 -0.0057 1.0000 0.0136 -6.750 -0.5923 0.07257 0.06894 -0.0097 1.0000 0.0136 -6.500 -0.5809 0.06577 0.06207 -0.0140 1.0000 0.0139 -6.250 -0.5688 0.05907 0.05524 -0.0179 1.0000 0.0145 -6.000 -0.5521 0.05391 0.04994 -0.0207 1.0000 0.0150 -5.750 -0.5320 0.04931 0.04517 -0.0232 1.0000 0.0155 -5.500 -0.5087 0.04485 0.04048 -0.0254 1.0000 0.0159 -5.250 -0.4820 0.04037 0.03570 -0.0273 1.0000 0.0158 -4.750 -0.4266 0.03240 0.02709 -0.0298 1.0000 0.0105 -4.500 -0.3964 0.02834 0.02260 -0.0309 1.0000 0.0092 -4.250 -0.3652 0.02463 0.01841 -0.0315 1.0000 0.0082 -4.000 -0.3336 0.02127 0.01454 -0.0317 1.0000 0.0074 -3.750 -0.3029 0.01849 0.01128 -0.0316 1.0000 0.0068 -3.500 -0.2733 0.01657 0.00874 -0.0312 1.0000 0.0058 -3.250 -0.2448 0.01478 0.00670 -0.0310 1.0000 0.0058 -3.000 -0.2162 0.01312 0.00486 -0.0311 1.0000 0.0061 -2.750 -0.1870 0.01185 0.00344 -0.0314 1.0000 0.0069 -2.500 -0.1584 0.01119 0.00268 -0.0316 1.0000 0.0108 -2.250 -0.1292 0.01049 0.00191 -0.0319 1.0000 0.0276 -2.000 -0.1006 0.00763 0.00163 -0.0337 1.0000 0.6812 -1.750 -0.0823 0.00729 0.00165 -0.0308 1.0000 0.8057 -1.500 -0.0674 0.00708 0.00159 -0.0271 1.0000 0.8789 -1.250 -0.0504 0.00687 0.00141 -0.0242 1.0000 0.9223 -1.000 -0.0263 0.00670 0.00118 -0.0231 1.0000 0.9635 -0.750 0.0010 0.00662 0.00106 -0.0231 1.0000 1.0000 -0.500 0.0296 0.00663 0.00102 -0.0235 1.0000 1.0000 -0.250 0.0579 0.00665 0.00102 -0.0238 1.0000 1.0000 0.000 0.0861 0.00667 0.00106 -0.0240 1.0000 1.0000 0.250 0.1140 0.00670 0.00114 -0.0242 1.0000 1.0000 0.500 0.1418 0.00674 0.00125 -0.0243 1.0000 1.0000 0.750 0.1695 0.00678 0.00144 -0.0244 1.0000 1.0000 1.000 0.1971 0.00683 0.00165 -0.0244 1.0000 1.0000 1.250 0.2495 0.01047 0.00195 -0.0298 0.0366 1.0000 1.500 0.2774 0.01126 0.00275 -0.0297 0.0111 1.0000 1.750 0.3053 0.01187 0.00343 -0.0297 0.0068 1.0000 2.000 0.3325 0.01308 0.00478 -0.0294 0.0059 1.0000 2.250 0.3596 0.01469 0.00655 -0.0289 0.0056 1.0000 2.500 0.3874 0.01649 0.00859 -0.0284 0.0056 1.0000 2.750 0.4163 0.01821 0.01086 -0.0278 0.0064 1.0000 3.000 0.4452 0.02073 0.01384 -0.0270 0.0069 1.0000 3.250 0.4742 0.02383 0.01745 -0.0261 0.0076 1.0000 3.500 0.5024 0.02724 0.02133 -0.0253 0.0085 1.0000 3.750 0.5296 0.03092 0.02544 -0.0247 0.0096 1.0000 4.000 0.5549 0.03497 0.02988 -0.0245 0.0111 1.0000 4.500 0.6052 0.04248 0.03800 -0.0243 0.0155 1.0000 4.750 0.6281 0.04667 0.04246 -0.0248 0.0153 1.0000 5.000 0.6486 0.05099 0.04699 -0.0255 0.0149 1.0000 5.250 0.6660 0.05573 0.05190 -0.0266 0.0144 1.0000 5.500 0.6795 0.06189 0.05824 -0.0284 0.0138 1.0000 5.750 0.6900 0.06935 0.06582 -0.0309 0.0133 1.0000 6.000 0.7062 0.07425 0.07085 -0.0330 0.0133 1.0000 6.250 0.7210 0.07921 0.07592 -0.0355 0.0133 1.0000 6.500 0.7345 0.08421 0.08102 -0.0385 0.0133 1.0000 6.750 0.7465 0.08924 0.08612 -0.0419 0.0133 1.0000 7.000 0.7570 0.09425 0.09117 -0.0457 0.0133 1.0000 7.250 0.7659 0.09923 0.09618 -0.0498 0.0132 1.0000 7.500 0.7730 0.10408 0.10104 -0.0541 0.0132 1.0000 7.750 0.7762 0.10856 0.10550 -0.0578 0.0132 1.0000 8.000 0.7791 0.11290 0.10982 -0.0608 0.0132 1.0000 8.250 0.7822 0.11716 0.11406 -0.0634 0.0132 1.0000