XFOIL Version 6.94 Calculated polar for: mta3 300.2.14.15.2.c13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2095 0.04350 0.03429 -0.0299 1.0000 0.2610 -2.750 -0.1725 0.04100 0.03134 -0.0324 1.0000 0.2416 -2.500 -0.1415 0.03853 0.02867 -0.0332 1.0000 0.2310 -2.250 -0.1075 0.03684 0.02651 -0.0339 1.0000 0.2208 -2.000 -0.0772 0.03492 0.02443 -0.0337 1.0000 0.2149 -1.750 -0.0446 0.03338 0.02263 -0.0335 1.0000 0.2080 -1.500 -0.0123 0.03230 0.02127 -0.0328 1.0000 0.2024 -1.250 0.0189 0.03107 0.02009 -0.0317 1.0000 0.2005 -1.000 0.0514 0.02985 0.01917 -0.0305 1.0000 0.2003 -0.500 0.1890 0.02976 0.01663 -0.0385 0.3608 0.2282 -0.250 0.2163 0.02981 0.01656 -0.0373 0.3408 0.2494 0.000 0.2447 0.02931 0.01665 -0.0367 0.3286 0.3184 0.250 0.2768 0.02823 0.01671 -0.0343 0.3202 1.0000 0.500 0.3111 0.02932 0.01706 -0.0335 0.3132 1.0000 0.750 0.3454 0.03059 0.01771 -0.0331 0.3080 1.0000 1.000 0.3800 0.03192 0.01864 -0.0332 0.3046 1.0000 1.250 0.4137 0.03315 0.01977 -0.0334 0.3011 1.0000 1.500 0.4464 0.03452 0.02110 -0.0337 0.2972 1.0000 1.750 0.4783 0.03604 0.02265 -0.0340 0.2947 1.0000 2.000 0.5096 0.03774 0.02446 -0.0343 0.2959 1.0000 2.250 0.5406 0.03969 0.02654 -0.0349 0.2991 1.0000 2.500 0.5701 0.04189 0.02885 -0.0354 0.3030 1.0000 2.750 0.5996 0.04402 0.03124 -0.0362 0.3090 1.0000 3.000 0.6283 0.04651 0.03413 -0.0377 0.3188 1.0000 3.250 0.6548 0.04946 0.03716 -0.0385 0.3253 1.0000 3.500 0.6804 0.05227 0.04046 -0.0410 0.3372 1.0000 3.750 0.7042 0.05578 0.04407 -0.0423 0.3465 1.0000 4.000 0.7258 0.05964 0.04835 -0.0462 0.3668 1.0000 4.250 0.7401 0.06462 0.05383 -0.0539 0.4022 1.0000 4.500 0.7448 0.07089 0.06038 -0.0635 0.4498 1.0000 4.750 0.7155 0.07845 0.06807 -0.0767 0.5208 1.0000 5.000 0.6699 0.08494 0.07452 -0.0858 0.6111 1.0000