XFOIL Version 6.96 Calculated polar for: GOE 419 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.3588 0.09910 0.09434 -0.0253 1.0000 0.0294 -7.750 -0.3601 0.09710 0.09242 -0.0255 1.0000 0.0298 -7.500 -0.3637 0.09534 0.09076 -0.0258 1.0000 0.0300 -7.250 -0.3618 0.09297 0.08846 -0.0275 1.0000 0.0302 -7.000 -0.3582 0.09050 0.08605 -0.0295 1.0000 0.0303 -6.750 -0.3529 0.08789 0.08349 -0.0316 1.0000 0.0305 -6.500 -0.3468 0.08507 0.08071 -0.0333 1.0000 0.0305 -6.250 -0.3396 0.08228 0.07793 -0.0349 1.0000 0.0306 -6.000 -0.3320 0.07940 0.07506 -0.0361 1.0000 0.0307 -5.750 -0.3238 0.07653 0.07218 -0.0371 1.0000 0.0307 -5.500 -0.3323 0.07214 0.06793 -0.0318 1.0000 0.0318 -5.250 -0.3255 0.06912 0.06492 -0.0312 0.9993 0.0334 -5.000 -0.2920 0.06465 0.06035 -0.0379 0.9940 0.0361 -4.750 -0.2495 0.06011 0.05563 -0.0466 0.9885 0.0387 -4.500 -0.1978 0.05610 0.05124 -0.0559 0.9822 0.0401 -4.250 -0.1653 0.05147 0.04640 -0.0599 0.9768 0.0400 -3.750 -0.1021 0.04013 0.03467 -0.0661 0.9660 0.0196 -3.500 -0.0668 0.03589 0.03011 -0.0693 0.9604 0.0191 -3.250 -0.0302 0.03182 0.02563 -0.0720 0.9554 0.0189 -3.000 0.0046 0.02803 0.02131 -0.0735 0.9494 0.0194 -2.750 0.0419 0.02452 0.01715 -0.0754 0.9455 0.0218 -2.500 0.0747 0.02211 0.01423 -0.0763 0.9398 0.0237 -2.250 0.1095 0.01984 0.01134 -0.0770 0.9348 0.0274 -2.000 0.1464 0.01818 0.00928 -0.0785 0.9311 0.0361 -1.750 0.1778 0.01711 0.00796 -0.0789 0.9240 0.0525 -1.500 0.2149 0.01615 0.00692 -0.0807 0.9185 0.0723 -1.250 0.2497 0.01548 0.00606 -0.0819 0.9095 0.0917 -1.000 0.2890 0.01474 0.00527 -0.0840 0.9021 0.1161 -0.750 0.3234 0.01419 0.00483 -0.0852 0.8921 0.1815 -0.500 0.3747 0.01205 0.00446 -0.0902 0.8883 1.0000 -0.250 0.4072 0.01205 0.00422 -0.0911 0.8779 1.0000 0.000 0.4391 0.01207 0.00405 -0.0918 0.8670 1.0000 0.250 0.4710 0.01209 0.00392 -0.0925 0.8561 1.0000 0.500 0.5023 0.01212 0.00383 -0.0931 0.8448 1.0000 0.750 0.5327 0.01217 0.00379 -0.0935 0.8329 1.0000 1.000 0.5620 0.01223 0.00380 -0.0937 0.8203 1.0000 1.250 0.5900 0.01231 0.00388 -0.0936 0.8067 1.0000 1.500 0.6170 0.01240 0.00396 -0.0933 0.7922 1.0000 1.750 0.6432 0.01250 0.00409 -0.0929 0.7771 1.0000 2.000 0.6690 0.01261 0.00425 -0.0924 0.7612 1.0000 2.250 0.6947 0.01273 0.00450 -0.0918 0.7447 1.0000 2.500 0.7206 0.01285 0.00472 -0.0912 0.7276 1.0000 2.750 0.7466 0.01298 0.00497 -0.0907 0.7104 1.0000 3.000 0.7685 0.01303 0.00486 -0.0884 0.6542 1.0000 3.250 0.7810 0.01364 0.00484 -0.0842 0.5169 1.0000 3.500 0.7806 0.01596 0.00531 -0.0792 0.2204 1.0000 3.750 0.7894 0.01841 0.00674 -0.0762 0.0292 1.0000 4.000 0.8100 0.01930 0.00788 -0.0746 0.0232 1.0000 4.250 0.8289 0.02037 0.00922 -0.0728 0.0200 1.0000 4.500 0.8472 0.02144 0.01055 -0.0709 0.0188 1.0000 4.750 0.8634 0.02267 0.01202 -0.0687 0.0179 1.0000 5.000 0.8780 0.02407 0.01353 -0.0663 0.0159 1.0000 5.250 0.8926 0.02591 0.01535 -0.0642 0.0142 1.0000 5.500 0.9147 0.02812 0.01752 -0.0630 0.0139 1.0000 5.750 0.9422 0.03043 0.01990 -0.0625 0.0139 1.0000 6.000 0.9703 0.03260 0.02225 -0.0618 0.0142 1.0000 6.250 0.9967 0.03449 0.02453 -0.0604 0.0153 1.0000 6.500 1.0221 0.03711 0.02762 -0.0589 0.0171 1.0000 6.750 1.0438 0.04019 0.03107 -0.0573 0.0185 1.0000 7.000 1.0619 0.04353 0.03475 -0.0556 0.0195 1.0000 7.250 1.0768 0.04725 0.03883 -0.0537 0.0202 1.0000