XFOIL Version 6.96 Calculated polar for: EPPLER 554 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 ------ -------- --------- --------- -------- -------- -------- -19.000 -0.7390 0.13829 0.13585 -0.0247 1.0000 0.0050 -18.750 -0.7574 0.13019 0.12763 -0.0288 1.0000 0.0049 -18.500 -0.7769 0.12215 0.11947 -0.0328 1.0000 0.0049 -18.250 -0.7964 0.11435 0.11153 -0.0367 1.0000 0.0050 -18.000 -0.8080 0.10829 0.10537 -0.0397 1.0000 0.0049 -17.750 -0.8194 0.10242 0.09939 -0.0426 1.0000 0.0049 -17.500 -0.8317 0.09660 0.09346 -0.0453 1.0000 0.0048 -17.250 -0.8418 0.09108 0.08783 -0.0480 1.0000 0.0051 -17.000 -0.8496 0.08646 0.08313 -0.0499 1.0000 0.0048 -16.750 -0.8556 0.08188 0.07844 -0.0520 1.0000 0.0051 -16.500 -0.8641 0.07733 0.07380 -0.0538 1.0000 0.0049 -16.250 -0.8680 0.07341 0.06978 -0.0555 1.0000 0.0050 -16.000 -0.8717 0.06963 0.06592 -0.0569 1.0000 0.0050 -15.750 -0.8744 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0.5195 3.750 0.8603 0.00892 0.00298 -0.0974 0.4267 0.5231 4.000 0.8876 0.00907 0.00311 -0.0974 0.4229 0.5268 4.250 0.9162 0.00910 0.00319 -0.0975 0.4208 0.5303 4.500 0.9446 0.00914 0.00327 -0.0977 0.4178 0.5336 4.750 0.9727 0.00918 0.00336 -0.0978 0.4144 0.5385 5.000 1.0002 0.00927 0.00346 -0.0978 0.4106 0.5431 5.250 1.0269 0.00943 0.00361 -0.0977 0.4064 0.5475 5.500 1.0553 0.00946 0.00370 -0.0978 0.4038 0.5516 5.750 1.0833 0.00950 0.00381 -0.0979 0.4005 0.5569 6.000 1.1107 0.00957 0.00392 -0.0979 0.3965 0.5622 6.250 1.1371 0.00972 0.00406 -0.0977 0.3917 0.5673 6.500 1.1646 0.00979 0.00419 -0.0977 0.3879 0.5733 6.750 1.1922 0.00984 0.00431 -0.0978 0.3836 0.5799 7.000 1.2188 0.00996 0.00445 -0.0977 0.3787 0.5862 7.250 1.2447 0.01010 0.00462 -0.0974 0.3734 0.5935 7.500 1.2721 0.01016 0.00476 -0.0974 0.3681 0.6011 7.750 1.2979 0.01029 0.00492 -0.0972 0.3612 0.6088 8.000 1.3239 0.01042 0.00509 -0.0970 0.3548 0.6178 8.250 1.3493 0.01057 0.00528 -0.0967 0.3464 0.6271 8.500 1.3743 0.01075 0.00548 -0.0963 0.3375 0.6377 8.750 1.3977 0.01098 0.00573 -0.0957 0.3261 0.6486 9.000 1.4203 0.01127 0.00601 -0.0949 0.3122 0.6599 9.250 1.4412 0.01162 0.00634 -0.0939 0.2948 0.6726 9.500 1.4608 0.01203 0.00672 -0.0927 0.2767 0.6863 9.750 1.4782 0.01252 0.00718 -0.0911 0.2571 0.7016 10.000 1.4918 0.01316 0.00774 -0.0890 0.2351 0.7189 10.250 1.5024 0.01376 0.00830 -0.0863 0.2135 0.7384 10.750 1.5157 0.01512 0.00962 -0.0796 0.1773 0.7894 11.000 1.5209 0.01581 0.01035 -0.0761 0.1623 0.8301 11.250 1.5223 0.01621 0.01098 -0.0718 0.1511 1.0000 11.500 1.5268 0.01712 0.01184 -0.0688 0.1382 1.0000 11.750 1.5308 0.01811 0.01280 -0.0659 0.1271 1.0000 12.000 1.5327 0.01927 0.01393 -0.0630 0.1167 1.0000 12.250 1.5330 0.02062 0.01526 -0.0602 0.1074 1.0000 12.500 1.5342 0.02208 0.01670 -0.0579 0.0990 1.0000 12.750 1.5339 0.02377 0.01840 -0.0558 0.0909 1.0000 13.000 1.5318 0.02576 0.02039 -0.0539 0.0839 1.0000 13.250 1.5301 0.02789 0.02252 -0.0524 0.0770 1.0000 13.500 1.5286 0.03016 0.02482 -0.0512 0.0713 1.0000 13.750 1.5233 0.03289 0.02755 -0.0501 0.0651 1.0000 14.000 1.5217 0.03543 0.03013 -0.0493 0.0604 1.0000 14.250 1.5157 0.03848 0.03319 -0.0487 0.0555 1.0000 14.500 1.5106 0.04155 0.03629 -0.0482 0.0508 1.0000 14.750 1.5059 0.04469 0.03946 -0.0480 0.0471 1.0000 15.000 1.5016 0.04788 0.04270 -0.0478 0.0437 1.0000 15.250 1.4956 0.05136 0.04621 -0.0479 0.0398 1.0000 15.500 1.4919 0.05471 0.04960 -0.0481 0.0373 1.0000 15.750 1.4876 0.05821 0.05315 -0.0485 0.0347 1.0000 16.000 1.4840 0.06173 0.05671 -0.0489 0.0322 1.0000 16.250 1.4794 0.06546 0.06049 -0.0495 0.0297 1.0000 16.500 1.4771 0.06897 0.06404 -0.0502 0.0277 1.0000 16.750 1.4734 0.07273 0.06785 -0.0510 0.0258 1.0000 17.000 1.4707 0.07647 0.07163 -0.0519 0.0239 1.0000 17.250 1.4665 0.08046 0.07567 -0.0530 0.0220 1.0000 17.500 1.4654 0.08407 0.07933 -0.0540 0.0206 1.0000