EPPLER 625 AIRFOIL (e625-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 625 AIRFOIL (e625-il) Reynolds number: 50,000 Max Cl/Cd: 18.78 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e625-il-50000.txt Download as CSV file: xf-e625-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 625 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.4097 0.12138 0.11468 -0.0047 1.0000 0.2241
-10.000 -0.3876 0.11641 0.10973 -0.0037 1.0000 0.2332
-9.750 -0.4162 0.11577 0.10924 -0.0064 1.0000 0.2417
-9.500 -0.3832 0.11021 0.10365 -0.0044 1.0000 0.2541
-9.250 -0.3815 0.10685 0.10037 -0.0047 1.0000 0.2636
-9.000 -0.4084 0.10585 0.09954 -0.0064 1.0000 0.2754
-8.750 -0.3881 0.10159 0.09530 -0.0050 1.0000 0.2910
-8.500 -0.3746 0.09789 0.09165 -0.0040 1.0000 0.3062
-8.250 -0.3742 0.09502 0.08887 -0.0033 1.0000 0.3237
-8.000 -0.3904 0.09298 0.08698 -0.0027 1.0000 0.3424
-7.750 -0.3521 0.08846 0.08244 -0.0005 1.0000 0.3656
-7.500 -0.3766 0.08730 0.08146 0.0015 1.0000 0.3924
-7.250 -0.3259 0.08305 0.07718 0.0045 1.0000 0.4339
-7.000 -0.3054 0.08052 0.07471 0.0078 1.0000 0.4816
-6.750 -0.2797 0.07807 0.07230 0.0110 1.0000 0.5365
-6.500 -0.2433 0.07496 0.06920 0.0132 1.0000 0.5971
-5.500 -0.3499 0.06410 0.05922 0.0160 1.0000 0.5144
-5.250 -0.4007 0.06369 0.05908 0.0234 1.0000 0.5140
-5.000 -0.4456 0.06356 0.05914 0.0309 1.0000 0.5167
-4.750 -0.4966 0.06263 0.05837 0.0372 1.0000 0.5140
-4.500 -0.5680 0.05251 0.04584 0.0180 1.0000 0.1795
-4.250 -0.5488 0.04915 0.04189 0.0185 0.9965 0.1526
-4.000 -0.5046 0.04548 0.03741 0.0151 0.9830 0.1391
-3.750 -0.4609 0.04231 0.03370 0.0118 0.9693 0.1338
-3.500 -0.4159 0.03991 0.03037 0.0094 0.9553 0.1277
-3.250 -0.3680 0.03762 0.02766 0.0061 0.9408 0.1278
-3.000 -0.3182 0.03567 0.02564 0.0020 0.9260 0.1349
-2.750 0.1031 0.02714 0.01962 -0.0546 0.9448 1.0000
-2.500 0.1828 0.02640 0.01835 -0.0646 0.9259 1.0000
-2.250 0.2420 0.02589 0.01747 -0.0706 0.9041 1.0000
-2.000 0.2843 0.02567 0.01697 -0.0734 0.8806 1.0000
-1.750 0.3184 0.02559 0.01665 -0.0742 0.8590 1.0000
-1.500 0.3386 0.02588 0.01677 -0.0729 0.8359 1.0000
-1.250 0.3596 0.02612 0.01684 -0.0715 0.8151 1.0000
-1.000 0.3803 0.02637 0.01692 -0.0699 0.7958 1.0000
-0.750 0.3994 0.02666 0.01706 -0.0680 0.7776 1.0000
-0.500 0.4175 0.02700 0.01725 -0.0660 0.7601 1.0000
-0.250 0.4348 0.02741 0.01755 -0.0640 0.7426 1.0000
0.000 0.4526 0.02781 0.01783 -0.0621 0.7262 1.0000
0.250 0.4706 0.02818 0.01809 -0.0601 0.7106 1.0000
0.500 0.4882 0.02866 0.01848 -0.0583 0.6948 1.0000
0.750 0.5058 0.02919 0.01894 -0.0566 0.6794 1.0000
1.000 0.5231 0.02976 0.01944 -0.0549 0.6643 1.0000
1.250 0.5404 0.03039 0.02003 -0.0532 0.6499 1.0000
1.500 0.5578 0.03104 0.02063 -0.0516 0.6363 1.0000
1.750 0.5763 0.03153 0.02103 -0.0498 0.6239 1.0000
2.000 0.5958 0.03190 0.02133 -0.0480 0.6122 1.0000
2.250 0.6108 0.03290 0.02235 -0.0466 0.5986 1.0000
2.500 0.6259 0.03387 0.02332 -0.0450 0.5861 1.0000
2.750 0.6452 0.03436 0.02374 -0.0432 0.5759 1.0000
3.000 0.6611 0.03523 0.02462 -0.0415 0.5645 1.0000
3.250 0.6719 0.03670 0.02613 -0.0400 0.5526 1.0000
3.500 0.6906 0.03728 0.02668 -0.0382 0.5433 1.0000
3.750 0.7025 0.03858 0.02800 -0.0364 0.5327 1.0000
4.000 0.7083 0.04051 0.03003 -0.0346 0.5222 1.0000
4.250 0.7331 0.04049 0.02992 -0.0328 0.5142 1.0000
4.500 0.7243 0.04376 0.03334 -0.0306 0.5034 1.0000
4.750 0.7491 0.04383 0.03335 -0.0289 0.4960 1.0000
5.000 0.7298 0.04781 0.03746 -0.0261 0.4859 1.0000
5.250 0.7398 0.04921 0.03887 -0.0239 0.4782 1.0000
5.500 0.7010 0.05451 0.04423 -0.0204 0.4702 1.0000
5.750 0.7282 0.05458 0.04431 -0.0187 0.4632 1.0000
6.000 0.6205 0.06409 0.05373 -0.0125 0.4597 1.0000
6.250 0.5811 0.06864 0.05822 -0.0092 0.4568 1.0000
6.500 0.5697 0.07182 0.06137 -0.0074 0.4535 1.0000
6.750 0.5848 0.07350 0.06304 -0.0061 0.4471 1.0000
7.000 0.5600 0.07779 0.06729 -0.0047 0.4463 1.0000
7.250 0.5489 0.08155 0.07104 -0.0040 0.4467 1.0000
7.500 0.5526 0.08498 0.07448 -0.0038 0.4484 1.0000
7.750 0.4396 0.09764 0.08716 -0.0093 0.5609 1.0000
8.000 0.4443 0.09930 0.08881 -0.0081 0.5481 1.0000
8.250 0.4402 0.10120 0.09069 -0.0067 0.5393 1.0000
8.500 0.4645 0.10451 0.09404 -0.0071 0.5308 1.0000
8.750 0.4541 0.10586 0.09536 -0.0055 0.5208 1.0000
9.000 0.4832 0.10992 0.09945 -0.0062 0.5128 1.0000
9.250 0.4741 0.11080 0.10032 -0.0046 0.5010 1.0000
9.500 0.4839 0.11393 0.10348 -0.0045 0.4950 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 625 AIRFOIL (e625-il)