XFOIL Version 6.94 Calculated polar for: lb3000 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0444 0.07718 0.05384 -0.0309 1.0001 0.9999 -2.750 -0.0624 0.07506 0.05195 -0.0292 1.0001 0.9999 -2.500 -0.0838 0.07279 0.04990 -0.0270 1.0001 0.9999 -2.250 -0.1053 0.07050 0.04767 -0.0249 1.0001 0.9999 -2.000 -0.1172 0.06864 0.04539 -0.0240 1.0001 0.9999 -1.750 -0.1119 0.06770 0.04352 -0.0248 1.0001 0.9999 -1.500 -0.0977 0.06747 0.04229 -0.0256 1.0001 0.9999 -1.250 -0.0811 0.06759 0.04149 -0.0260 1.0001 0.9999 -1.000 -0.0638 0.06789 0.04100 -0.0262 1.0001 0.9999 -0.750 -0.0463 0.06833 0.04070 -0.0262 1.0001 0.9999 -0.500 -0.0288 0.06886 0.04061 -0.0262 1.0001 0.9999 -0.250 -0.0113 0.06948 0.04067 -0.0261 1.0001 0.9999 0.000 0.0061 0.07018 0.04087 -0.0260 1.0001 0.9999 0.250 0.0233 0.07096 0.04119 -0.0259 1.0001 0.9999 0.500 0.0403 0.07183 0.04166 -0.0258 1.0001 0.9999 0.750 0.0571 0.07278 0.04226 -0.0257 1.0001 0.9999 1.000 0.0735 0.07381 0.04299 -0.0257 1.0001 0.9999 1.250 0.0895 0.07494 0.04384 -0.0257 1.0001 0.9999 1.500 0.1051 0.07618 0.04484 -0.0258 1.0001 0.9999 1.750 0.1202 0.07752 0.04599 -0.0259 1.0001 0.9999 2.000 0.1346 0.07898 0.04728 -0.0261 1.0001 0.9999 2.250 0.1484 0.08057 0.04872 -0.0264 1.0001 0.9999 2.500 0.1614 0.08229 0.05033 -0.0268 1.0001 0.9999 2.750 0.1736 0.08416 0.05209 -0.0273 1.0001 0.9999 3.000 0.1850 0.08617 0.05402 -0.0278 1.0001 0.9999 3.250 0.1957 0.08832 0.05609 -0.0285 1.0001 0.9999 3.500 0.2057 0.09059 0.05829 -0.0292 1.0001 0.9999 3.750 0.2154 0.09299 0.06060 -0.0301 1.0001 0.9999 4.000 0.2249 0.09548 0.06302 -0.0310 1.0001 0.9999 4.250 0.2343 0.09803 0.06549 -0.0320 1.0001 0.9999 4.500 0.2437 0.10063 0.06800 -0.0331 1.0001 0.9999 4.750 0.2533 0.10325 0.07054 -0.0342 1.0001 0.9999 5.000 0.2631 0.10589 0.07309 -0.0353 1.0001 0.9999