XFOIL Version 6.96 Calculated polar for: NACA 64-208 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.5514 0.09410 0.08926 -0.0148 1.0000 0.0933 -8.500 -0.5698 0.08979 0.08508 -0.0225 1.0000 0.0949 -8.250 -0.5887 0.08553 0.08078 -0.0300 1.0000 0.0954 -8.000 -0.5543 0.08197 0.07731 -0.0203 1.0000 0.1030 -7.750 -0.5638 0.07713 0.07251 -0.0265 1.0000 0.1067 -7.500 -0.5888 0.07390 0.06886 -0.0357 1.0000 0.1094 -7.250 -0.5097 0.05761 0.05330 -0.0309 1.0000 0.1257 -7.000 -0.5056 0.05384 0.04951 -0.0311 1.0000 0.1340 -6.750 -0.5525 0.06099 0.05619 -0.0336 1.0000 0.1331 -6.500 -0.5450 0.05725 0.05240 -0.0340 1.0000 0.1435 -6.250 -0.5363 0.05391 0.04896 -0.0342 1.0000 0.1580 -6.000 -0.5286 0.05089 0.04588 -0.0339 1.0000 0.1822 -5.750 -0.5186 0.04843 0.04347 -0.0323 1.0000 0.2104 -5.000 -0.4394 0.03159 0.02388 -0.0353 1.0000 0.0841 -4.750 -0.4141 0.02831 0.01988 -0.0336 1.0000 0.0700 -4.500 -0.3910 0.02613 0.01740 -0.0323 1.0000 0.0671 -4.250 -0.3682 0.02447 0.01535 -0.0310 1.0000 0.0682 -4.000 -0.3463 0.02201 0.01276 -0.0300 1.0000 0.0711 -3.750 -0.3235 0.02036 0.01099 -0.0287 1.0000 0.0720 -3.500 -0.3014 0.01899 0.00954 -0.0273 1.0000 0.0748 -3.250 -0.2798 0.01792 0.00842 -0.0260 1.0000 0.0801 -3.000 -0.2594 0.01679 0.00743 -0.0248 1.0000 0.0910 -2.750 -0.2386 0.01578 0.00647 -0.0237 1.0000 0.1032 -2.500 -0.2165 0.01462 0.00554 -0.0231 1.0000 0.1460 -2.250 -0.2154 0.01212 0.00584 -0.0166 1.0000 0.7530 -2.000 -0.2138 0.01229 0.00612 -0.0097 1.0000 0.8352 -1.750 -0.2138 0.01237 0.00627 -0.0026 1.0000 0.8997 -1.500 -0.0751 0.01282 0.00622 -0.0192 1.0000 0.9971 -1.250 -0.0864 0.01258 0.00596 -0.0150 1.0000 1.0000 -1.000 -0.1013 0.01234 0.00566 -0.0101 1.0000 1.0000 -0.750 -0.0884 0.01235 0.00553 -0.0094 1.0000 1.0000 -0.500 -0.0685 0.01250 0.00553 -0.0096 1.0000 1.0000 -0.250 -0.0270 0.01286 0.00575 -0.0136 0.9934 1.0000 0.000 0.0187 0.01323 0.00601 -0.0182 0.9845 1.0000 0.250 0.0618 0.01357 0.00625 -0.0223 0.9752 1.0000 0.500 0.1071 0.01392 0.00655 -0.0266 0.9670 1.0000 0.750 0.1491 0.01422 0.00683 -0.0302 0.9576 1.0000 1.000 0.1888 0.01450 0.00712 -0.0332 0.9476 1.0000 1.250 0.2340 0.01478 0.00743 -0.0372 0.9399 1.0000 1.500 0.2731 0.01502 0.00773 -0.0400 0.9299 1.0000 1.750 0.3110 0.01527 0.00808 -0.0425 0.9199 1.0000 2.000 0.3599 0.01541 0.00836 -0.0468 0.9127 1.0000 2.250 0.3978 0.01556 0.00864 -0.0488 0.9012 1.0000 2.500 0.4336 0.01569 0.00893 -0.0503 0.8891 1.0000 2.750 0.4695 0.01575 0.00921 -0.0515 0.8762 1.0000 3.000 0.5122 0.01449 0.00812 -0.0501 0.8428 1.0000 3.250 0.5314 0.01322 0.00680 -0.0436 0.7875 1.0000 3.500 0.5518 0.01288 0.00654 -0.0402 0.7440 1.0000 3.750 0.5695 0.01257 0.00613 -0.0358 0.6611 1.0000 4.000 0.5627 0.01584 0.00646 -0.0288 0.1424 1.0000 4.250 0.5808 0.01759 0.00786 -0.0273 0.0992 1.0000 4.500 0.6011 0.01907 0.00923 -0.0259 0.0839 1.0000 4.750 0.6250 0.02046 0.01064 -0.0249 0.0764 1.0000 5.000 0.6504 0.02254 0.01257 -0.0243 0.0706 1.0000 5.250 0.6775 0.02394 0.01420 -0.0237 0.0647 1.0000 5.500 0.7056 0.02606 0.01648 -0.0231 0.0629 1.0000 5.750 0.7333 0.02855 0.01930 -0.0223 0.0630 1.0000 6.000 0.7598 0.03164 0.02291 -0.0212 0.0656 1.0000 6.250 0.7834 0.03476 0.02650 -0.0201 0.0664 1.0000 6.500 0.8049 0.03830 0.03050 -0.0189 0.0677 1.0000 10.750 0.6923 0.12340 0.11921 -0.0312 0.1113 1.0000 11.000 0.6655 0.12821 0.12396 -0.0373 0.1101 1.0000 11.250 0.6525 0.13234 0.12806 -0.0410 0.1055 1.0000