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 = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.6578 0.09432 0.09274 0.0200 1.0000 0.0047 -7.750 -0.6536 0.09021 0.08864 0.0170 1.0000 0.0047 -7.500 -0.6439 0.08532 0.08376 0.0115 1.0000 0.0047 -7.250 -0.6306 0.07983 0.07826 0.0042 1.0000 0.0047 -7.000 -0.6152 0.07422 0.07262 -0.0024 1.0000 0.0047 -6.750 -0.5984 0.06877 0.06711 -0.0080 1.0000 0.0048 -6.500 -0.5800 0.06346 0.06172 -0.0127 1.0000 0.0048 -6.250 -0.5602 0.05830 0.05646 -0.0167 1.0000 0.0048 -6.000 -0.5389 0.05330 0.05134 -0.0199 1.0000 0.0048 -5.750 -0.5164 0.04848 0.04637 -0.0226 1.0000 0.0048 -5.500 -0.4924 0.04385 0.04158 -0.0247 1.0000 0.0048 -5.250 -0.4679 0.03563 0.03304 -0.0282 1.0000 0.0051 -5.000 -0.4414 0.03008 0.02717 -0.0304 1.0000 0.0054 -4.750 -0.4136 0.02639 0.02321 -0.0317 1.0000 0.0058 -4.500 -0.3853 0.02351 0.02009 -0.0327 1.0000 0.0062 -4.250 -0.3559 0.02086 0.01718 -0.0334 1.0000 0.0068 -4.000 -0.3258 0.01843 0.01448 -0.0339 1.0000 0.0076 -3.750 -0.2955 0.01634 0.01203 -0.0342 1.0000 0.0087 -3.500 -0.2656 0.01475 0.01023 -0.0342 1.0000 0.0103 -3.000 -0.2029 0.00950 0.00438 -0.0341 1.0000 0.0055 -2.750 -0.1740 0.00873 0.00355 -0.0341 1.0000 0.0046 -2.500 -0.1430 0.00744 0.00210 -0.0347 1.0000 0.0046 -2.250 -0.1127 0.00668 0.00116 -0.0352 1.0000 0.0060 -2.000 -0.0840 0.00633 0.00082 -0.0353 1.0000 0.0279 -1.750 -0.0535 0.00531 0.00064 -0.0367 1.0000 0.2855 -1.500 -0.0230 0.00396 0.00062 -0.0383 1.0000 0.6914 -1.250 0.0040 0.00381 0.00061 -0.0381 1.0000 0.7443 -1.000 0.0303 0.00368 0.00065 -0.0377 1.0000 0.8000 -0.750 0.0563 0.00359 0.00069 -0.0372 1.0000 0.8354 -0.500 0.0821 0.00354 0.00072 -0.0367 1.0000 0.8589 -0.250 0.1081 0.00350 0.00075 -0.0363 1.0000 0.8777 0.000 0.1336 0.00346 0.00080 -0.0357 1.0000 0.8953 0.250 0.1587 0.00341 0.00085 -0.0350 1.0000 0.9133 0.750 0.2175 0.00652 0.00101 -0.0365 0.0661 0.9456 1.000 0.2417 0.00676 0.00110 -0.0355 0.0096 1.0000 1.250 0.2706 0.00731 0.00178 -0.0356 0.0047 1.0000 1.500 0.2990 0.00806 0.00265 -0.0355 0.0042 1.0000 1.750 0.3263 0.00969 0.00443 -0.0350 0.0044 1.0000 2.000 0.3551 0.01018 0.00497 -0.0348 0.0056 1.0000 9.250 0.8113 0.13275 0.13135 -0.0803 0.0044 1.0000 9.500 0.8145 0.13692 0.13552 -0.0820 0.0044 1.0000 9.750 0.8179 0.14101 0.13960 -0.0837 0.0044 1.0000