XFOIL Version 6.96 Calculated polar for: WORTMANN FX 66-17A-175 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -12.750 -0.1273 0.11012 0.10653 -0.0687 0.6253 0.0185 -12.500 -0.1262 0.10566 0.10208 -0.0706 0.6211 0.0185 -12.250 -0.1385 0.09836 0.09481 -0.0731 0.6193 0.0192 -12.000 -0.1356 0.09494 0.09135 -0.0742 0.6147 0.0195 -11.750 -0.1306 0.09213 0.08851 -0.0751 0.6104 0.0200 -11.500 -0.1299 0.08795 0.08433 -0.0767 0.6066 0.0202 -11.250 -0.1270 0.08463 0.08100 -0.0779 0.6026 0.0210 -11.000 -0.1310 0.07936 0.07572 -0.0800 0.5999 0.0214 -10.750 -0.1363 0.07387 0.07022 -0.0823 0.5975 0.0222 -10.500 -0.1580 0.06322 0.05961 -0.0874 0.5971 0.0215 -9.000 -0.3741 0.03454 0.02944 -0.1069 0.5943 0.0122 -8.750 -0.3752 0.03080 0.02537 -0.1055 0.5905 0.0116 -8.500 -0.3665 0.02808 0.02228 -0.1042 0.5866 0.0117 -8.250 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0.01175 0.00443 -0.0968 0.4796 0.5417 1.500 0.6512 0.01179 0.00448 -0.0970 0.4776 0.5435 1.750 0.6793 0.01184 0.00452 -0.0972 0.4757 0.5452 2.000 0.7074 0.01190 0.00457 -0.0974 0.4738 0.5470 2.250 0.7354 0.01193 0.00459 -0.0976 0.4720 0.5489 2.500 0.7634 0.01199 0.00466 -0.0978 0.4703 0.5509 2.750 0.7918 0.01215 0.00480 -0.0981 0.4684 0.5528 3.000 0.8194 0.01223 0.00492 -0.0983 0.4670 0.5547 3.250 0.8469 0.01229 0.00505 -0.0984 0.4654 0.5568 3.500 0.8745 0.01238 0.00517 -0.0985 0.4637 0.5590 3.750 0.9022 0.01248 0.00529 -0.0987 0.4621 0.5611 4.000 0.9298 0.01256 0.00541 -0.0988 0.4605 0.5634 4.250 0.9574 0.01264 0.00554 -0.0990 0.4590 0.5660 4.500 0.9851 0.01275 0.00568 -0.0992 0.4576 0.5688 4.750 1.0130 0.01288 0.00584 -0.0994 0.4562 0.5720 5.000 1.0412 0.01310 0.00607 -0.0998 0.4547 0.5749 5.250 1.0686 0.01329 0.00631 -0.1000 0.4534 0.5776 5.500 1.0950 0.01338 0.00652 -0.0999 0.4521 0.5806 5.750 1.1215 0.01351 0.00675 -0.0999 0.4507 0.5842 6.000 1.1482 0.01366 0.00699 -0.1000 0.4492 0.5882 6.250 1.1750 0.01380 0.00722 -0.1001 0.4477 0.5922 6.500 1.2016 0.01392 0.00743 -0.1001 0.4461 0.5964 6.750 1.2286 0.01404 0.00763 -0.1002 0.4444 0.6014 7.000 1.2561 0.01408 0.00764 -0.1003 0.4402 0.6068 7.250 1.2759 0.01375 0.00749 -0.0989 0.4328 0.6125 7.500 1.3001 0.01363 0.00732 -0.0983 0.4253 0.6199 7.750 1.3208 0.01344 0.00736 -0.0971 0.4188 0.6282 8.000 1.3424 0.01333 0.00730 -0.0961 0.4104 0.6389 8.250 1.3613 0.01318 0.00727 -0.0945 0.3955 0.6532 8.500 1.3577 0.01389 0.00764 -0.0892 0.3157 0.6713 8.750 1.3309 0.01578 0.00916 -0.0806 0.2367 0.6978 9.000 1.2956 0.01773 0.01115 -0.0712 0.1770 0.8165 9.250 1.2693 0.02071 0.01404 -0.0661 0.1195 1.0000 9.500 1.2475 0.02383 0.01707 -0.0625 0.0891 1.0000 9.750 1.2158 0.02877 0.02189 -0.0602 0.0528 1.0000 10.000 1.1913 0.03360 0.02665 -0.0589 0.0255 1.0000 10.250 1.1744 0.03777 0.03083 -0.0579 0.0139 1.0000 10.500 1.1687 0.04091 0.03404 -0.0571 0.0115 1.0000 10.750 1.1658 0.04377 0.03699 -0.0565 0.0098 1.0000 11.000 1.1634 0.04661 0.03989 -0.0560 0.0091 1.0000 11.250 1.1592 0.04968 0.04304 -0.0555 0.0085 1.0000 11.500 1.1523 0.05313 0.04659 -0.0551 0.0079 1.0000 11.750 1.1494 0.05629 0.04983 -0.0548 0.0077 1.0000 12.000 1.1518 0.05892 0.05253 -0.0547 0.0075 1.0000 12.250 1.1501 0.06209 0.05578 -0.0546 0.0074 1.0000 12.500 1.1499 0.06513 0.05889 -0.0546 0.0072 1.0000 12.750 1.1492 0.06831 0.06214 -0.0547 0.0068 1.0000 13.000 1.1491 0.07142 0.06532 -0.0548 0.0067 1.0000 13.250 1.1486 0.07465 0.06861 -0.0549 0.0064 1.0000 13.500 1.1480 0.07793 0.07195 -0.0552 0.0060 1.0000 13.750 1.1473 0.08127 0.07536 -0.0554 0.0060 1.0000 14.000 1.1477 0.08445 0.07861 -0.0557 0.0058 1.0000 14.250 1.1454 0.08804 0.08225 -0.0561 0.0056 1.0000 14.500 1.1505 0.09068 0.08496 -0.0564 0.0059 1.0000