XFOIL Version 6.96 Calculated polar for: EPPLER 904 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 ------ -------- --------- --------- -------- -------- -------- -10.000 -0.3783 0.09347 0.09136 -0.0376 1.0000 0.0103 -9.750 -0.3825 0.08948 0.08740 -0.0380 1.0000 0.0104 -8.000 -0.5153 0.06922 0.06715 -0.0494 0.9939 0.0090 -7.750 -0.5051 0.06295 0.06077 -0.0553 0.9875 0.0093 -7.500 -0.4911 0.05722 0.05488 -0.0602 0.9836 0.0095 -7.250 -0.4806 0.05221 0.04970 -0.0623 0.9772 0.0099 -7.000 -0.4654 0.04726 0.04454 -0.0644 0.9727 0.0104 -6.750 -0.4459 0.04247 0.03950 -0.0662 0.9702 0.0112 -6.500 -0.4323 0.03846 0.03522 -0.0654 0.9638 0.0120 -6.250 -0.4106 0.03471 0.03117 -0.0656 0.9607 0.0131 -6.000 -0.3773 0.03395 0.03011 -0.0658 0.9590 0.0151 -5.000 -0.2992 0.01912 0.01375 -0.0618 0.9455 0.0130 -4.750 -0.2684 0.01577 0.00987 -0.0608 0.9441 0.0075 -4.500 -0.2447 0.01463 0.00854 -0.0597 0.9404 0.0072 -4.250 -0.2219 0.01278 0.00651 -0.0585 0.9373 0.0076 -4.000 -0.1986 0.01155 0.00517 -0.0576 0.9345 0.0093 -3.750 -0.1709 0.01112 0.00470 -0.0578 0.9326 0.0131 -3.500 -0.1443 0.01053 0.00397 -0.0576 0.9307 0.0139 -3.250 -0.1227 0.01021 0.00352 -0.0562 0.9271 0.0151 -3.000 -0.1081 0.00889 0.00302 -0.0540 0.9232 0.2161 -2.750 -0.1045 0.00710 0.00277 -0.0500 0.9194 0.6077 -2.500 -0.0932 0.00668 0.00315 -0.0458 0.9165 0.8383 -2.250 -0.0712 0.00681 0.00319 -0.0444 0.9134 0.8618 -2.000 -0.0468 0.00694 0.00326 -0.0435 0.9108 0.8775 -1.750 -0.0210 0.00703 0.00330 -0.0429 0.9087 0.8888 -1.500 0.0070 0.00714 0.00336 -0.0427 0.9070 0.8978 -1.250 0.0350 0.00723 0.00341 -0.0426 0.9053 0.9064 -1.000 0.0589 0.00735 0.00347 -0.0417 0.9026 0.9146 -0.750 0.0851 0.00756 0.00367 -0.0409 0.8997 0.9244 -0.500 0.1180 0.00780 0.00390 -0.0415 0.8975 0.9345 -0.250 0.1488 0.00795 0.00403 -0.0419 0.8950 0.9431 0.000 0.2003 0.00809 0.00415 -0.0468 0.8936 0.9458 0.250 0.2491 0.00813 0.00416 -0.0514 0.8911 0.9477 0.500 0.2829 0.00815 0.00419 -0.0527 0.8852 0.9512 0.750 0.3074 0.00811 0.00412 -0.0518 0.8790 0.9561 1.000 0.3442 0.00806 0.00409 -0.0539 0.8725 0.9574 1.250 0.3799 0.00801 0.00406 -0.0557 0.8677 0.9588 1.500 0.4141 0.00799 0.00411 -0.0572 0.8639 0.9602 1.750 0.4437 0.00797 0.00416 -0.0578 0.8583 0.9623 2.000 0.4737 0.00793 0.00416 -0.0583 0.8533 0.9644 2.250 0.4955 0.00793 0.00423 -0.0572 0.8470 0.9678 2.500 0.5199 0.00775 0.00406 -0.0561 0.8291 0.9698 2.750 0.5260 0.00792 0.00337 -0.0501 0.6048 0.9733 3.000 0.5218 0.00969 0.00380 -0.0443 0.3134 0.9785 3.250 0.5201 0.01182 0.00464 -0.0390 0.0113 0.9831 3.500 0.5486 0.01225 0.00522 -0.0393 0.0105 0.9844 3.750 0.5746 0.01288 0.00598 -0.0390 0.0104 0.9860 4.000 0.5978 0.01376 0.00698 -0.0382 0.0105 0.9880 4.250 0.6183 0.01497 0.00830 -0.0367 0.0109 0.9900 4.500 0.6409 0.01689 0.01028 -0.0356 0.0120 0.9913 5.000 0.6926 0.01813 0.01171 -0.0349 0.0091 0.9942 5.250 0.7216 0.02027 0.01411 -0.0345 0.0093 0.9946 6.250 0.8024 0.03670 0.03198 -0.0284 0.0116 1.0000 6.500 0.8035 0.03877 0.03429 -0.0236 0.0116 1.0000 6.750 0.8027 0.04047 0.03621 -0.0185 0.0115 1.0000 7.000 0.8085 0.04033 0.03626 -0.0138 0.0109 1.0000 7.250 0.8133 0.04110 0.03718 -0.0090 0.0099 1.0000 7.500 0.8124 0.04301 0.03927 -0.0039 0.0093 1.0000 7.750 0.8123 0.04554 0.04200 0.0005 0.0089 1.0000 8.000 0.8126 0.04840 0.04506 0.0042 0.0084 1.0000 8.250 0.8112 0.05143 0.04828 0.0077 0.0080 1.0000 8.500 0.8063 0.05471 0.05174 0.0113 0.0077 1.0000 8.750 0.7995 0.05784 0.05503 0.0145 0.0076 1.0000 9.000 0.7870 0.06066 0.05798 0.0184 0.0075 1.0000 9.250 0.7713 0.06363 0.06105 0.0219 0.0075 1.0000 9.500 0.7547 0.06696 0.06449 0.0241 0.0075 1.0000 9.750 0.7375 0.07078 0.06841 0.0249 0.0075 1.0000 10.000 0.7187 0.07555 0.07326 0.0240 0.0076 1.0000 10.250 0.7007 0.08154 0.07932 0.0209 0.0079 1.0000