XFOIL Version 6.96 Calculated polar for: EPPLER E850 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 ------ -------- --------- --------- -------- -------- -------- -9.000 -0.4439 0.08243 0.08037 -0.0314 1.0000 0.0124 -8.750 -0.4509 0.07837 0.07634 -0.0315 1.0000 0.0125 -8.500 -0.4579 0.07485 0.07285 -0.0312 1.0000 0.0128 -8.250 -0.4684 0.07064 0.06867 -0.0314 1.0000 0.0127 -8.000 -0.4816 0.06628 0.06434 -0.0318 1.0000 0.0127 -7.750 -0.5046 0.05978 0.05786 -0.0350 1.0000 0.0122 -7.500 -0.5307 0.05718 0.05523 -0.0345 1.0000 0.0118 -7.250 -0.5313 0.05034 0.04824 -0.0419 0.9967 0.0119 -7.000 -0.5258 0.04384 0.04156 -0.0472 0.9940 0.0122 -6.750 -0.5178 0.03902 0.03660 -0.0499 0.9909 0.0127 -6.500 -0.5056 0.03468 0.03209 -0.0521 0.9879 0.0134 -6.250 -0.4890 0.03041 0.02761 -0.0543 0.9854 0.0143 -6.000 -0.4679 0.02629 0.02324 -0.0566 0.9836 0.0155 -5.750 -0.4432 0.02226 0.01890 -0.0586 0.9822 0.0172 -5.500 -0.4142 0.01903 0.01532 -0.0600 0.9811 0.0199 -5.250 -0.3903 0.01998 0.01595 -0.0583 0.9773 0.0220 -5.000 -0.3733 0.01315 0.00853 -0.0594 0.9751 0.0238 -4.750 -0.3474 0.01059 0.00582 -0.0602 0.9736 0.0260 -4.500 -0.3186 0.00871 0.00371 -0.0607 0.9723 0.0284 -4.250 -0.3122 0.01860 0.01305 -0.0588 0.9737 0.0195 -4.000 -0.2815 0.01581 0.00994 -0.0578 0.9731 0.0096 -3.750 -0.2524 0.01387 0.00781 -0.0577 0.9722 0.0080 -3.500 -0.2214 0.01259 0.00635 -0.0583 0.9712 0.0070 -3.250 -0.1886 0.01182 0.00539 -0.0593 0.9702 0.0065 -3.000 -0.1550 0.01144 0.00484 -0.0605 0.9693 0.0065 -2.750 -0.1229 0.01125 0.00448 -0.0614 0.9679 0.0076 -2.500 -0.1004 0.01109 0.00421 -0.0603 0.9640 0.0093 -2.250 -0.0724 0.00975 0.00388 -0.0615 0.9622 0.2791 -2.000 -0.0430 0.00906 0.00390 -0.0626 0.9605 0.4811 -1.750 -0.0121 0.00886 0.00412 -0.0634 0.9589 0.5970 -1.500 0.0213 0.00885 0.00416 -0.0647 0.9576 0.6261 -1.250 0.0559 0.00882 0.00416 -0.0663 0.9565 0.6391 -1.000 0.0829 0.00878 0.00413 -0.0662 0.9531 0.6489 -0.750 0.1117 0.00871 0.00409 -0.0665 0.9495 0.6588 -0.500 0.1459 0.00861 0.00402 -0.0680 0.9473 0.6694 -0.250 0.1820 0.00850 0.00395 -0.0698 0.9455 0.6807 0.000 0.2202 0.00835 0.00386 -0.0721 0.9441 0.6927 0.250 0.2610 0.00815 0.00373 -0.0749 0.9430 0.7054 0.500 0.2851 0.00802 0.00368 -0.0740 0.9363 0.7187 0.750 0.3268 0.00771 0.00348 -0.0769 0.9339 0.7332 1.000 0.3751 0.00726 0.00315 -0.0810 0.9317 0.7486 1.250 0.4116 0.00685 0.00286 -0.0824 0.9238 0.7656 1.500 0.4507 0.00637 0.00258 -0.0842 0.9124 0.7840 1.750 0.4820 0.00611 0.00246 -0.0846 0.9002 0.8044 2.000 0.5097 0.00589 0.00237 -0.0840 0.8812 0.8277 2.250 0.5350 0.00571 0.00226 -0.0829 0.8435 0.8551 2.500 0.5575 0.00572 0.00217 -0.0810 0.7692 0.8891 2.750 0.5631 0.00660 0.00232 -0.0757 0.5790 0.9482 3.000 0.5738 0.00838 0.00283 -0.0730 0.2924 1.0000 3.250 0.5880 0.00979 0.00335 -0.0709 0.0947 1.0000 3.500 0.6058 0.01117 0.00444 -0.0688 0.0132 1.0000 3.750 0.6279 0.01208 0.00549 -0.0673 0.0113 1.0000 4.000 0.6495 0.01314 0.00665 -0.0658 0.0104 1.0000 4.250 0.6731 0.01382 0.00734 -0.0651 0.0069 1.0000 4.500 0.6940 0.01680 0.01056 -0.0633 0.0057 1.0000 4.750 0.7201 0.01750 0.01137 -0.0628 0.0051 1.0000 5.000 0.7460 0.01954 0.01366 -0.0618 0.0047 1.0000 5.250 0.7698 0.02258 0.01707 -0.0602 0.0047 1.0000 5.500 0.7878 0.02692 0.02191 -0.0575 0.0052 1.0000 5.750 0.8125 0.03169 0.02718 -0.0539 0.0097 1.0000 6.000 0.8276 0.03545 0.03129 -0.0511 0.0097 1.0000 6.250 0.8410 0.03903 0.03518 -0.0484 0.0094 1.0000 6.500 0.8523 0.04276 0.03919 -0.0458 0.0090 1.0000 6.750 0.8617 0.04644 0.04313 -0.0433 0.0085 1.0000 7.000 0.8691 0.05016 0.04707 -0.0409 0.0082 1.0000 7.250 0.8741 0.05394 0.05106 -0.0387 0.0079 1.0000 7.500 0.8766 0.05774 0.05503 -0.0366 0.0076 1.0000 7.750 0.8765 0.06160 0.05906 -0.0345 0.0074 1.0000 8.000 0.8731 0.06556 0.06316 -0.0326 0.0072 1.0000 8.250 0.8670 0.06941 0.06713 -0.0308 0.0071 1.0000 8.500 0.8546 0.07307 0.07089 -0.0285 0.0071 1.0000 8.750 0.8388 0.07703 0.07493 -0.0268 0.0071 1.0000 9.000 0.8225 0.08179 0.07977 -0.0272 0.0072 1.0000