EPPLER 904 AIRFOIL (e904-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 904 AIRFOIL (e904-il) Reynolds number: 200,000 Max Cl/Cd: 53.87 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e904-il-200000-n5.txt Download as CSV file: xf-e904-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4167 0.07307 0.06997 -0.0396 1.0000 0.0086
-8.250 -0.4223 0.06818 0.06510 -0.0424 0.9980 0.0086
-8.000 -0.4280 0.06118 0.05809 -0.0495 0.9938 0.0085
-7.750 -0.4359 0.05552 0.05236 -0.0538 0.9868 0.0084
-7.250 -0.4949 0.05867 0.05505 -0.0537 0.9789 0.0085
-7.000 -0.4886 0.05361 0.04980 -0.0551 0.9718 0.0080
-6.750 -0.4797 0.04884 0.04478 -0.0557 0.9657 0.0075
-6.500 -0.4685 0.04439 0.04006 -0.0556 0.9599 0.0070
-6.250 -0.4522 0.03961 0.03490 -0.0557 0.9565 0.0065
-6.000 -0.4409 0.03562 0.03054 -0.0537 0.9504 0.0060
-5.750 -0.4228 0.03140 0.02582 -0.0524 0.9468 0.0054
-5.500 -0.3997 0.02756 0.02138 -0.0515 0.9444 0.0049
-5.250 -0.3770 0.02474 0.01797 -0.0502 0.9414 0.0046
-5.000 -0.3551 0.02270 0.01553 -0.0488 0.9377 0.0044
-4.750 -0.3289 0.02083 0.01330 -0.0482 0.9355 0.0043
-4.500 -0.3013 0.01924 0.01145 -0.0480 0.9337 0.0042
-4.250 -0.2737 0.01776 0.00977 -0.0479 0.9322 0.0043
-4.000 -0.2461 0.01653 0.00839 -0.0480 0.9308 0.0043
-3.750 -0.2258 0.01564 0.00735 -0.0465 0.9270 0.0046
-3.500 -0.2030 0.01492 0.00644 -0.0455 0.9238 0.0051
-3.250 -0.1770 0.01439 0.00573 -0.0451 0.9217 0.0060
-3.000 -0.1494 0.01387 0.00510 -0.0451 0.9198 0.0137
-2.750 -0.1426 0.01123 0.00464 -0.0425 0.9171 0.4524
-2.500 -0.1381 0.01031 0.00494 -0.0369 0.9124 0.7546
-2.250 -0.1112 0.01051 0.00511 -0.0360 0.9102 0.8331
-2.000 -0.0849 0.01062 0.00508 -0.0354 0.9081 0.8593
-1.750 -0.0562 0.01072 0.00508 -0.0354 0.9064 0.8784
-1.500 -0.0243 0.01089 0.00516 -0.0359 0.9053 0.8959
-1.250 0.0098 0.01116 0.00536 -0.0366 0.9045 0.9168
-1.000 0.0587 0.01151 0.00559 -0.0405 0.9048 0.9320
-0.750 0.1014 0.01165 0.00567 -0.0436 0.9044 0.9407
-0.500 0.1448 0.01172 0.00569 -0.0472 0.9035 0.9447
-0.250 0.1802 0.01180 0.00574 -0.0491 0.9009 0.9488
0.000 0.2069 0.01184 0.00577 -0.0491 0.8974 0.9543
0.250 0.2426 0.01180 0.00573 -0.0511 0.8951 0.9565
0.500 0.2799 0.01173 0.00566 -0.0534 0.8929 0.9581
0.750 0.3151 0.01166 0.00561 -0.0552 0.8900 0.9599
1.000 0.3432 0.01164 0.00564 -0.0555 0.8838 0.9629
1.250 0.3781 0.01147 0.00550 -0.0569 0.8787 0.9644
1.500 0.4010 0.01140 0.00548 -0.0559 0.8694 0.9679
1.750 0.4329 0.01123 0.00550 -0.0567 0.8602 0.9693
2.000 0.4664 0.01108 0.00544 -0.0578 0.8525 0.9704
2.250 0.4946 0.01105 0.00554 -0.0580 0.8441 0.9725
2.500 0.5248 0.01097 0.00559 -0.0586 0.8367 0.9742
2.750 0.5521 0.01093 0.00571 -0.0585 0.8276 0.9764
3.000 0.5764 0.01070 0.00561 -0.0572 0.8012 0.9788
3.250 0.5737 0.01142 0.00471 -0.0491 0.4617 0.9829
3.500 0.5658 0.01433 0.00552 -0.0437 0.0629 0.9882
3.750 0.5849 0.01544 0.00659 -0.0421 0.0126 0.9913
4.000 0.6071 0.01612 0.00747 -0.0410 0.0108 0.9937
4.250 0.6299 0.01707 0.00858 -0.0401 0.0097 0.9960
4.500 0.6489 0.01871 0.01030 -0.0388 0.0064 0.9985
4.750 0.6731 0.01925 0.01093 -0.0383 0.0049 1.0000
5.000 0.6880 0.02036 0.01215 -0.0357 0.0042 1.0000
5.250 0.7058 0.02179 0.01372 -0.0335 0.0038 1.0000
5.500 0.7257 0.02348 0.01561 -0.0317 0.0037 1.0000
5.750 0.7451 0.02541 0.01782 -0.0296 0.0035 1.0000
6.000 0.7619 0.02759 0.02032 -0.0270 0.0035 1.0000
6.250 0.7750 0.02989 0.02297 -0.0238 0.0036 1.0000
6.500 0.7842 0.03241 0.02585 -0.0198 0.0037 1.0000
6.750 0.7903 0.03490 0.02867 -0.0156 0.0038 1.0000
7.000 0.7940 0.03759 0.03167 -0.0110 0.0039 1.0000
7.250 0.7957 0.04044 0.03483 -0.0064 0.0041 1.0000
7.500 0.7968 0.04333 0.03811 -0.0021 0.0042 1.0000
7.750 0.7965 0.04638 0.04143 0.0019 0.0044 1.0000
8.000 0.7946 0.04955 0.04484 0.0058 0.0045 1.0000
8.250 0.7906 0.05272 0.04824 0.0095 0.0047 1.0000
8.500 0.7824 0.05609 0.05182 0.0133 0.0048 1.0000
8.750 0.7710 0.05907 0.05496 0.0171 0.0050 1.0000
9.000 0.7568 0.06200 0.05802 0.0208 0.0050 1.0000
9.250 0.7425 0.06509 0.06124 0.0233 0.0051 1.0000
9.500 0.7256 0.06880 0.06506 0.0247 0.0052 1.0000
9.750 0.7111 0.07278 0.06915 0.0247 0.0052 1.0000
10.000 0.6959 0.07751 0.07397 0.0232 0.0052 1.0000
10.250 0.6822 0.08344 0.07996 0.0198 0.0052 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 904 AIRFOIL (e904-il)