EPPLER 325 AIRFOIL (e325-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 325 AIRFOIL (e325-il) Reynolds number: 1,000,000 Max Cl/Cd: 82.6 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e325-il-1000000-n5.txt Download as CSV file: xf-e325-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 325 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.9416 0.03779 0.03496 -0.0060 1.0000 0.0026
-10.750 -0.9724 0.03195 0.02857 0.0013 1.0000 0.0026
-10.500 -0.9629 0.03134 0.02793 0.0040 1.0000 0.0027
-10.250 -0.9545 0.03019 0.02666 0.0068 1.0000 0.0027
-10.000 -0.9528 0.02738 0.02352 0.0105 1.0000 0.0027
-9.750 -0.9401 0.02612 0.02212 0.0127 1.0000 0.0028
-9.500 -0.9260 0.02493 0.02077 0.0147 1.0000 0.0029
-9.250 -0.9129 0.02322 0.01884 0.0170 1.0000 0.0029
-9.000 -0.8962 0.02210 0.01757 0.0187 1.0000 0.0030
-8.750 -0.8574 0.02087 0.01617 0.0157 0.9453 0.0031
-8.500 -0.8303 0.01984 0.01464 0.0156 0.8308 0.0032
-8.250 -0.8118 0.01899 0.01350 0.0173 0.7870 0.0032
-8.000 -0.7912 0.01831 0.01262 0.0185 0.7554 0.0033
-7.750 -0.7698 0.01749 0.01159 0.0197 0.7298 0.0034
-7.500 -0.7478 0.01680 0.01071 0.0207 0.7055 0.0035
-7.250 -0.7251 0.01624 0.01000 0.0216 0.6846 0.0036
-7.000 -0.7022 0.01571 0.00935 0.0225 0.6665 0.0036
-6.750 -0.6794 0.01514 0.00864 0.0234 0.6483 0.0037
-6.500 -0.6564 0.01464 0.00803 0.0243 0.6327 0.0037
-6.250 -0.6338 0.01411 0.00739 0.0253 0.6172 0.0037
-6.000 -0.6103 0.01377 0.00695 0.0261 0.6027 0.0038
-5.750 -0.5876 0.01332 0.00641 0.0270 0.5894 0.0039
-5.500 -0.5644 0.01296 0.00598 0.0279 0.5769 0.0039
-5.250 -0.5417 0.01255 0.00549 0.0289 0.5643 0.0042
-5.000 -0.5187 0.01220 0.00507 0.0298 0.5525 0.0044
-4.750 -0.4953 0.01191 0.00472 0.0307 0.5409 0.0045
-4.500 -0.4714 0.01166 0.00442 0.0315 0.5304 0.0047
-4.250 -0.4477 0.01139 0.00409 0.0323 0.5206 0.0049
-4.000 -0.4239 0.01115 0.00380 0.0332 0.5095 0.0052
-3.750 -0.3996 0.01096 0.00356 0.0339 0.4992 0.0052
-3.500 -0.3752 0.01076 0.00330 0.0346 0.4906 0.0056
-3.000 -0.3264 0.01041 0.00286 0.0360 0.4718 0.0080
-2.750 -0.3019 0.01024 0.00268 0.0367 0.4629 0.0117
-2.500 -0.2775 0.01005 0.00251 0.0374 0.4549 0.0213
-2.250 -0.2533 0.00988 0.00237 0.0381 0.4465 0.0364
-2.000 -0.2302 0.00961 0.00222 0.0390 0.4379 0.0702
-1.750 -0.2083 0.00929 0.00208 0.0401 0.4310 0.1272
-1.500 -0.1935 0.00856 0.00186 0.0424 0.4243 0.2752
-1.250 -0.2042 0.00670 0.00135 0.0496 0.4196 0.6274
-1.000 -0.1952 0.00607 0.00132 0.0538 0.4137 0.7867
-0.750 -0.1697 0.00614 0.00140 0.0545 0.4059 0.8152
-0.500 -0.1429 0.00623 0.00146 0.0549 0.3996 0.8287
-0.250 -0.1155 0.00630 0.00148 0.0550 0.3927 0.8356
0.000 -0.0880 0.00641 0.00155 0.0553 0.3855 0.8431
0.250 -0.0609 0.00650 0.00159 0.0555 0.3796 0.8496
0.500 -0.0325 0.00655 0.00161 0.0554 0.3734 0.8523
0.750 -0.0042 0.00659 0.00160 0.0553 0.3675 0.8532
1.000 0.0242 0.00663 0.00159 0.0551 0.3613 0.8542
1.250 0.0525 0.00669 0.00159 0.0550 0.3555 0.8550
1.500 0.0809 0.00672 0.00159 0.0548 0.3508 0.8558
1.750 0.1092 0.00677 0.00161 0.0547 0.3449 0.8568
2.000 0.1373 0.00684 0.00163 0.0545 0.3391 0.8579
2.250 0.1657 0.00688 0.00165 0.0544 0.3348 0.8588
2.500 0.1939 0.00694 0.00167 0.0542 0.3293 0.8597
2.750 0.2219 0.00701 0.00171 0.0541 0.3238 0.8606
3.000 0.2502 0.00706 0.00174 0.0539 0.3193 0.8614
3.250 0.2783 0.00712 0.00178 0.0537 0.3143 0.8622
3.500 0.3062 0.00721 0.00183 0.0536 0.3087 0.8630
3.750 0.3344 0.00727 0.00188 0.0534 0.3041 0.8638
4.000 0.3623 0.00735 0.00195 0.0533 0.2985 0.8647
4.250 0.3902 0.00743 0.00201 0.0531 0.2927 0.8657
4.500 0.4182 0.00750 0.00208 0.0530 0.2870 0.8666
4.750 0.4458 0.00760 0.00217 0.0529 0.2808 0.8675
5.000 0.4736 0.00768 0.00226 0.0527 0.2754 0.8684
5.250 0.5013 0.00778 0.00235 0.0526 0.2692 0.8693
5.500 0.5286 0.00790 0.00247 0.0525 0.2630 0.8703
5.750 0.5562 0.00800 0.00257 0.0524 0.2567 0.8712
6.000 0.5832 0.00815 0.00270 0.0523 0.2494 0.8722
6.250 0.6106 0.00826 0.00283 0.0522 0.2428 0.8733
6.500 0.6373 0.00843 0.00297 0.0522 0.2347 0.8743
6.750 0.6642 0.00857 0.00311 0.0522 0.2261 0.8754
7.000 0.6906 0.00876 0.00328 0.0522 0.2166 0.8766
7.250 0.7167 0.00896 0.00346 0.0522 0.2070 0.8778
7.500 0.7430 0.00915 0.00364 0.0522 0.1976 0.8790
7.750 0.7687 0.00937 0.00385 0.0523 0.1874 0.8800
8.000 0.7940 0.00963 0.00408 0.0525 0.1758 0.8813
8.250 0.8186 0.00991 0.00434 0.0527 0.1630 0.8826
8.500 0.8427 0.01023 0.00463 0.0530 0.1495 0.8840
8.750 0.8661 0.01060 0.00495 0.0534 0.1341 0.8854
9.000 0.8890 0.01098 0.00530 0.0539 0.1201 0.8869
9.250 0.9112 0.01141 0.00568 0.0544 0.1057 0.8886
9.500 0.9320 0.01191 0.00612 0.0552 0.0896 0.8904
9.750 0.9517 0.01247 0.00661 0.0560 0.0739 0.8923
10.000 0.9733 0.01288 0.00701 0.0567 0.0662 0.8941
10.250 0.9926 0.01342 0.00750 0.0576 0.0557 0.8958
10.500 1.0108 0.01396 0.00803 0.0587 0.0464 0.8982
10.750 1.0262 0.01460 0.00864 0.0601 0.0353 0.9011
11.000 1.0437 0.01509 0.00915 0.0613 0.0312 0.9040
11.250 1.0586 0.01568 0.00974 0.0629 0.0256 0.9073
11.500 1.0734 0.01622 0.01030 0.0644 0.0220 0.9105
11.750 1.0830 0.01679 0.01091 0.0668 0.0183 0.9146
12.000 1.0899 0.01736 0.01154 0.0697 0.0163 0.9202
12.250 1.0961 0.01803 0.01226 0.0724 0.0144 0.9268
12.500 1.1017 0.01877 0.01309 0.0749 0.0132 0.9354
12.750 1.1082 0.01966 0.01407 0.0768 0.0117 0.9466
13.000 1.1201 0.02089 0.01539 0.0767 0.0103 0.9602
13.500 1.1632 0.02425 0.01891 0.0707 0.0074 0.9794
13.750 1.1845 0.02625 0.02099 0.0672 0.0069 0.9864
14.250 1.1996 0.03099 0.02588 0.0640 0.0057 1.0000
14.500 1.1990 0.03349 0.02847 0.0637 0.0056 1.0000
14.750 1.1952 0.03644 0.03149 0.0632 0.0050 1.0000
15.000 1.1938 0.03921 0.03435 0.0626 0.0051 1.0000
15.250 1.1859 0.04274 0.03795 0.0617 0.0044 1.0000
15.500 1.1783 0.04630 0.04158 0.0608 0.0036 1.0000
15.750 1.1737 0.04958 0.04494 0.0599 0.0038 1.0000
16.000 1.1668 0.05313 0.04859 0.0588 0.0040 1.0000
16.250 1.1592 0.05689 0.05244 0.0577 0.0040 1.0000
16.500 1.1516 0.06077 0.05640 0.0564 0.0041 1.0000
16.750 1.1427 0.06494 0.06065 0.0548 0.0037 1.0000
17.000 1.1334 0.06922 0.06501 0.0533 0.0035 1.0000
17.250 1.1248 0.07347 0.06934 0.0517 0.0034 1.0000
17.500 1.1144 0.07807 0.07402 0.0499 0.0030 1.0000
17.750 1.1063 0.08241 0.07844 0.0482 0.0031 1.0000
18.000 1.0977 0.08687 0.08298 0.0464 0.0029 1.0000
18.250 1.0889 0.09145 0.08765 0.0445 0.0030 1.0000
18.500 1.0804 0.09606 0.09233 0.0426 0.0030 1.0000
18.750 1.0703 0.10094 0.09730 0.0405 0.0028 1.0000
19.000 1.0618 0.10569 0.10212 0.0385 0.0027 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 325 AIRFOIL (e325-il)