EPPLER 338 AIRFOIL (e338-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER 338 AIRFOIL (e338-il) Reynolds number: 1,000,000 Max Cl/Cd: 108.24 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e338-il-1000000-n5.txt Download as CSV file: xf-e338-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 338 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.2737 0.09332 0.09035 -0.0460 0.6446 0.0061
-9.750 -0.2728 0.08968 0.08668 -0.0475 0.6334 0.0061
-9.250 -0.2771 0.08123 0.07819 -0.0515 0.6134 0.0060
-8.750 -0.2290 0.06337 0.06040 -0.0484 0.5696 0.0055
-8.500 -0.2363 0.05886 0.05590 -0.0510 0.5641 0.0054
-8.250 -0.2534 0.05364 0.05067 -0.0553 0.5602 0.0053
-7.750 -0.2898 0.04608 0.04302 -0.0544 0.5511 0.0051
-6.750 -0.4224 0.01959 0.01412 -0.0339 0.5572 0.0036
-6.500 -0.4041 0.01742 0.01155 -0.0321 0.5487 0.0036
-6.250 -0.3816 0.01629 0.01017 -0.0310 0.5398 0.0037
-6.000 -0.3579 0.01544 0.00914 -0.0302 0.5317 0.0037
-5.750 -0.3340 0.01468 0.00820 -0.0294 0.5227 0.0037
-5.500 -0.3113 0.01354 0.00685 -0.0284 0.5144 0.0041
-5.250 -0.2864 0.01313 0.00634 -0.0277 0.5056 0.0042
-5.000 -0.2607 0.01287 0.00603 -0.0273 0.4975 0.0044
-4.750 -0.2353 0.01256 0.00564 -0.0268 0.4904 0.0046
-4.500 -0.2100 0.01222 0.00523 -0.0262 0.4836 0.0047
-4.250 -0.1848 0.01190 0.00484 -0.0256 0.4764 0.0049
-4.000 -0.1597 0.01156 0.00443 -0.0250 0.4698 0.0051
-3.750 -0.1342 0.01134 0.00414 -0.0244 0.4624 0.0053
-3.500 -0.1087 0.01110 0.00385 -0.0239 0.4562 0.0054
-3.250 -0.0831 0.01088 0.00357 -0.0233 0.4499 0.0056
-2.750 -0.0323 0.01043 0.00304 -0.0222 0.4389 0.0060
-2.500 -0.0065 0.01027 0.00284 -0.0217 0.4328 0.0062
-2.250 0.0194 0.01014 0.00268 -0.0213 0.4270 0.0067
-2.000 0.0456 0.01002 0.00253 -0.0209 0.4218 0.0074
-1.750 0.0717 0.00990 0.00237 -0.0205 0.4167 0.0081
-1.500 0.0977 0.00979 0.00224 -0.0200 0.4119 0.0096
-1.250 0.1241 0.00967 0.00211 -0.0197 0.4078 0.0118
-1.000 0.1504 0.00958 0.00202 -0.0193 0.4030 0.0156
-0.750 0.1765 0.00950 0.00194 -0.0189 0.3979 0.0249
-0.500 0.2024 0.00936 0.00188 -0.0185 0.3940 0.0482
-0.250 0.2278 0.00919 0.00184 -0.0181 0.3903 0.0912
0.000 0.2522 0.00894 0.00179 -0.0175 0.3863 0.1686
0.250 0.2406 0.00662 0.00170 -0.0099 0.3838 0.8350
0.500 0.2651 0.00678 0.00190 -0.0089 0.3798 0.8791
0.750 0.2913 0.00692 0.00200 -0.0083 0.3764 0.8922
1.000 0.3176 0.00705 0.00211 -0.0078 0.3729 0.9003
1.250 0.3435 0.00718 0.00220 -0.0072 0.3693 0.9070
1.500 0.3688 0.00736 0.00235 -0.0064 0.3655 0.9134
1.750 0.3961 0.00743 0.00238 -0.0062 0.3627 0.9160
2.000 0.4242 0.00747 0.00239 -0.0063 0.3594 0.9168
2.250 0.4522 0.00752 0.00240 -0.0064 0.3560 0.9175
2.500 0.4799 0.00759 0.00242 -0.0065 0.3525 0.9183
2.750 0.5077 0.00766 0.00245 -0.0065 0.3491 0.9191
3.000 0.5358 0.00771 0.00248 -0.0067 0.3465 0.9196
3.250 0.5637 0.00776 0.00251 -0.0068 0.3434 0.9202
3.500 0.5912 0.00782 0.00255 -0.0068 0.3398 0.9209
3.750 0.6185 0.00791 0.00261 -0.0068 0.3364 0.9215
4.000 0.6460 0.00797 0.00266 -0.0069 0.3335 0.9221
4.250 0.6736 0.00804 0.00272 -0.0069 0.3305 0.9227
4.500 0.7011 0.00811 0.00280 -0.0070 0.3273 0.9234
4.750 0.7281 0.00821 0.00287 -0.0070 0.3236 0.9242
5.000 0.7548 0.00832 0.00297 -0.0069 0.3198 0.9250
5.250 0.7823 0.00839 0.00305 -0.0070 0.3169 0.9257
5.500 0.8095 0.00847 0.00314 -0.0070 0.3130 0.9264
5.750 0.8363 0.00858 0.00324 -0.0070 0.3092 0.9271
6.000 0.8628 0.00871 0.00335 -0.0069 0.3045 0.9279
6.250 0.8898 0.00880 0.00345 -0.0070 0.2994 0.9287
6.500 0.9157 0.00896 0.00358 -0.0069 0.2919 0.9295
6.750 0.9422 0.00908 0.00370 -0.0068 0.2857 0.9303
7.000 0.9679 0.00925 0.00384 -0.0067 0.2781 0.9311
7.250 0.9940 0.00938 0.00399 -0.0066 0.2728 0.9319
7.500 1.0199 0.00953 0.00414 -0.0065 0.2672 0.9327
7.750 1.0452 0.00972 0.00431 -0.0064 0.2618 0.9335
8.000 1.0705 0.00989 0.00449 -0.0062 0.2542 0.9344
8.250 1.0946 0.01012 0.00469 -0.0058 0.2452 0.9354
8.500 1.1182 0.01034 0.00491 -0.0054 0.2364 0.9365
8.750 1.1416 0.01058 0.00514 -0.0049 0.2274 0.9375
9.000 1.1640 0.01086 0.00540 -0.0042 0.2181 0.9387
9.500 1.2073 0.01147 0.00598 -0.0027 0.1978 0.9412
9.750 1.2279 0.01181 0.00630 -0.0019 0.1881 0.9426
10.000 1.2476 0.01218 0.00666 -0.0008 0.1784 0.9442
10.250 1.2662 0.01258 0.00704 0.0004 0.1678 0.9460
10.500 1.2843 0.01299 0.00743 0.0016 0.1577 0.9478
10.750 1.2995 0.01346 0.00787 0.0033 0.1464 0.9498
11.000 1.3100 0.01393 0.00833 0.0060 0.1355 0.9528
11.250 1.3203 0.01442 0.00882 0.0086 0.1268 0.9565
11.500 1.3307 0.01495 0.00934 0.0111 0.1182 0.9611
11.750 1.3368 0.01572 0.01007 0.0140 0.1054 0.9672
12.250 1.3607 0.01765 0.01196 0.0161 0.0821 0.9820
12.500 1.3784 0.01875 0.01306 0.0155 0.0721 0.9914
12.750 1.3754 0.02048 0.01473 0.0177 0.0577 1.0000
13.000 1.3756 0.02211 0.01636 0.0194 0.0514 1.0000
13.250 1.3787 0.02371 0.01799 0.0205 0.0482 1.0000
13.500 1.3811 0.02551 0.01985 0.0213 0.0458 1.0000
13.750 1.3815 0.02763 0.02201 0.0219 0.0419 1.0000
14.000 1.3841 0.02967 0.02412 0.0223 0.0404 1.0000
14.250 1.3771 0.03269 0.02717 0.0226 0.0361 1.0000
14.500 1.3775 0.03512 0.02966 0.0226 0.0339 1.0000
14.750 1.3815 0.03723 0.03187 0.0226 0.0355 1.0000
15.000 1.3823 0.03970 0.03442 0.0225 0.0354 1.0000
15.250 1.3779 0.04272 0.03752 0.0223 0.0329 1.0000
15.500 1.3657 0.04664 0.04146 0.0220 0.0292 1.0000
15.750 1.3634 0.04957 0.04448 0.0216 0.0294 1.0000
16.000 1.3496 0.05390 0.04882 0.0209 0.0243 1.0000
16.250 1.3462 0.05714 0.05214 0.0203 0.0242 1.0000
16.750 1.3269 0.06534 0.06045 0.0184 0.0189 1.0000
17.000 1.3169 0.06963 0.06478 0.0172 0.0167 1.0000
17.250 1.3106 0.07353 0.06876 0.0161 0.0160 1.0000
17.500 1.3079 0.07704 0.07234 0.0151 0.0160 1.0000
17.750 1.2996 0.08132 0.07669 0.0137 0.0142 1.0000
18.000 1.2904 0.08580 0.08122 0.0122 0.0126 1.0000
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