EPPLER 904 AIRFOIL (e904-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER 904 AIRFOIL (e904-il) Reynolds number: 100,000 Max Cl/Cd: 41.95 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e904-il-100000-n5.txt Download as CSV file: xf-e904-il-100000-n5.csv |
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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3731 0.11083 0.10595 -0.0363 1.0000 0.0166
-10.500 -0.3746 0.10715 0.10231 -0.0369 1.0000 0.0158
-10.250 -0.4626 0.11046 0.10527 -0.0345 1.0000 0.0176
-10.000 -0.4632 0.10659 0.10146 -0.0352 1.0000 0.0165
-9.750 -0.4639 0.10294 0.09786 -0.0361 1.0000 0.0156
-9.500 -0.4663 0.09895 0.09394 -0.0371 1.0000 0.0149
-9.250 -0.4690 0.09519 0.09023 -0.0381 1.0000 0.0143
-9.000 -0.4728 0.09141 0.08652 -0.0392 1.0000 0.0138
-8.750 -0.4795 0.08724 0.08243 -0.0405 1.0000 0.0133
-8.500 -0.4883 0.08315 0.07843 -0.0419 1.0000 0.0128
-8.250 -0.4999 0.07960 0.07496 -0.0431 1.0000 0.0127
-8.000 -0.5185 0.07657 0.07200 -0.0420 1.0000 0.0125
-7.750 -0.5386 0.07431 0.06979 -0.0389 1.0000 0.0125
-7.500 -0.5554 0.07164 0.06712 -0.0361 1.0000 0.0123
-7.250 -0.5705 0.06871 0.06416 -0.0331 1.0000 0.0119
-7.000 -0.5826 0.06533 0.06070 -0.0303 1.0000 0.0115
-6.750 -0.5874 0.06148 0.05669 -0.0279 0.9997 0.0108
-6.500 -0.5687 0.05751 0.05245 -0.0290 0.9956 0.0102
-6.250 -0.5530 0.05421 0.04885 -0.0294 0.9919 0.0097
-6.000 -0.5447 0.04963 0.04398 -0.0292 0.9875 0.0095
-5.750 -0.5323 0.04532 0.03931 -0.0290 0.9838 0.0092
-5.500 -0.5171 0.04146 0.03503 -0.0285 0.9804 0.0090
-5.250 -0.5025 0.03803 0.03116 -0.0271 0.9764 0.0088
-5.000 -0.4843 0.03485 0.02749 -0.0259 0.9733 0.0087
-4.750 -0.4621 0.03184 0.02395 -0.0250 0.9709 0.0086
-4.500 -0.4365 0.02923 0.02068 -0.0246 0.9690 0.0085
-4.250 -0.4149 0.02713 0.01816 -0.0231 0.9663 0.0086
-4.000 -0.3912 0.02535 0.01605 -0.0221 0.9639 0.0086
-3.750 -0.3657 0.02374 0.01416 -0.0214 0.9616 0.0089
-3.500 -0.3396 0.02253 0.01273 -0.0209 0.9595 0.0097
-3.250 -0.3151 0.02109 0.01112 -0.0205 0.9576 0.0130
-3.000 -0.2884 0.02043 0.01028 -0.0204 0.9556 0.0202
-2.750 -0.2687 0.01946 0.00916 -0.0186 0.9522 0.0281
-2.500 -0.2617 0.01626 0.00868 -0.0155 0.9497 0.5053
-2.250 -0.2304 0.01635 0.00982 -0.0126 0.9500 0.8910
-2.000 -0.1862 0.01677 0.00974 -0.0152 0.9496 0.9266
-1.750 -0.1000 0.01741 0.00995 -0.0264 0.9550 0.9598
-1.500 -0.0074 0.01759 0.00981 -0.0397 0.9608 0.9910
-1.250 0.0419 0.01754 0.00958 -0.0449 0.9610 1.0000
-1.000 0.0698 0.01755 0.00941 -0.0457 0.9578 1.0000
-0.750 0.0939 0.01756 0.00933 -0.0457 0.9538 1.0000
-0.500 0.1161 0.01758 0.00928 -0.0453 0.9490 1.0000
-0.250 0.1442 0.01761 0.00926 -0.0460 0.9455 1.0000
0.000 0.1713 0.01766 0.00927 -0.0465 0.9418 1.0000
0.250 0.1917 0.01771 0.00930 -0.0456 0.9361 1.0000
0.500 0.2218 0.01775 0.00935 -0.0466 0.9324 1.0000
0.750 0.2470 0.01781 0.00942 -0.0467 0.9275 1.0000
1.000 0.2711 0.01786 0.00950 -0.0464 0.9218 1.0000
1.250 0.3059 0.01788 0.00958 -0.0483 0.9181 1.0000
1.500 0.3263 0.01793 0.00970 -0.0472 0.9105 1.0000
1.750 0.3638 0.01788 0.00976 -0.0494 0.9058 1.0000
2.000 0.3888 0.01787 0.01002 -0.0490 0.8974 1.0000
2.250 0.4370 0.01761 0.00996 -0.0530 0.8920 1.0000
2.500 0.4685 0.01741 0.00995 -0.0536 0.8815 1.0000
2.750 0.5019 0.01718 0.00994 -0.0544 0.8713 1.0000
3.000 0.5388 0.01691 0.00996 -0.0559 0.8624 1.0000
3.250 0.5733 0.01665 0.01003 -0.0567 0.8525 1.0000
3.500 0.6141 0.01464 0.00678 -0.0522 0.4676 1.0000
3.750 0.5973 0.01771 0.00751 -0.0446 0.0747 1.0000
4.000 0.6060 0.01912 0.00856 -0.0409 0.0249 1.0000
4.250 0.6189 0.01996 0.00959 -0.0378 0.0192 1.0000
4.500 0.6292 0.02096 0.01077 -0.0343 0.0172 1.0000
4.750 0.6367 0.02230 0.01226 -0.0304 0.0159 1.0000
5.000 0.6492 0.02373 0.01376 -0.0273 0.0152 1.0000
5.250 0.6688 0.02535 0.01549 -0.0254 0.0146 1.0000
5.500 0.6934 0.02737 0.01771 -0.0242 0.0142 1.0000
5.750 0.7179 0.02966 0.02029 -0.0229 0.0136 1.0000
6.000 0.7368 0.03142 0.02239 -0.0207 0.0116 1.0000
6.250 0.7521 0.03319 0.02446 -0.0182 0.0100 1.0000
6.500 0.7600 0.03469 0.02614 -0.0154 0.0076 1.0000
6.750 0.7684 0.03737 0.02916 -0.0123 0.0072 1.0000
7.000 0.7759 0.04007 0.03222 -0.0088 0.0071 1.0000
7.250 0.7810 0.04289 0.03540 -0.0051 0.0071 1.0000
7.500 0.7861 0.04546 0.03845 -0.0016 0.0072 1.0000
7.750 0.7900 0.04812 0.04144 0.0019 0.0074 1.0000
8.000 0.7902 0.05103 0.04467 0.0055 0.0074 1.0000
8.250 0.7891 0.05404 0.04797 0.0091 0.0077 1.0000
8.500 0.7834 0.05732 0.05152 0.0127 0.0078 1.0000
8.750 0.7742 0.06056 0.05499 0.0164 0.0081 1.0000
9.000 0.7606 0.06367 0.05828 0.0201 0.0083 1.0000
9.250 0.7463 0.06708 0.06186 0.0227 0.0085 1.0000
9.500 0.7301 0.07085 0.06576 0.0241 0.0086 1.0000
9.750 0.7147 0.07502 0.07004 0.0242 0.0088 1.0000
10.000 0.6994 0.07993 0.07504 0.0227 0.0090 1.0000
10.250 0.6857 0.08572 0.08089 0.0195 0.0091 1.0000
10.500 0.6763 0.09280 0.08799 0.0145 0.0094 1.0000
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