EPPLER 558 AIRFOIL (e558-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 558 AIRFOIL (e558-il) Reynolds number: 500,000 Max Cl/Cd: 99.07 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e558-il-500000-n5.txt Download as CSV file: xf-e558-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 558 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.250 -0.7436 0.09759 0.09370 -0.0495 1.0000 0.0089
-16.000 -0.7492 0.09322 0.08927 -0.0514 1.0000 0.0089
-15.750 -0.7541 0.08904 0.08503 -0.0532 1.0000 0.0089
-15.500 -0.7601 0.08478 0.08072 -0.0550 1.0000 0.0089
-15.250 -0.7657 0.08068 0.07656 -0.0567 1.0000 0.0090
-15.000 -0.7727 0.07639 0.07220 -0.0585 1.0000 0.0091
-14.750 -0.7800 0.07223 0.06798 -0.0603 1.0000 0.0091
-14.500 -0.7875 0.06811 0.06381 -0.0619 1.0000 0.0091
-14.250 -0.7967 0.06395 0.05957 -0.0636 1.0000 0.0093
-14.000 -0.8067 0.05997 0.05552 -0.0650 1.0000 0.0092
-13.750 -0.8180 0.05594 0.05143 -0.0664 1.0000 0.0093
-13.500 -0.8299 0.05200 0.04741 -0.0676 1.0000 0.0094
-13.250 -0.8438 0.04799 0.04334 -0.0687 1.0000 0.0094
-13.000 -0.8571 0.04424 0.03951 -0.0695 1.0000 0.0094
-12.750 -0.8713 0.04058 0.03579 -0.0700 1.0000 0.0094
-12.500 -0.8881 0.03694 0.03207 -0.0701 1.0000 0.0095
-12.250 -0.9100 0.03338 0.02843 -0.0695 1.0000 0.0094
-12.000 -0.8953 0.02918 0.02407 -0.0770 0.9956 0.0096
-11.750 -0.8696 0.02627 0.02099 -0.0833 0.9912 0.0098
-11.500 -0.8449 0.02427 0.01889 -0.0867 0.9865 0.0100
-11.250 -0.8161 0.02265 0.01717 -0.0899 0.9833 0.0103
-11.000 -0.7938 0.02148 0.01592 -0.0908 0.9773 0.0106
-10.750 -0.7650 0.02045 0.01482 -0.0925 0.9728 0.0109
-10.500 -0.7416 0.01955 0.01385 -0.0928 0.9659 0.0114
-9.750 -0.6642 0.01716 0.01127 -0.0943 0.9415 0.0131
-9.500 -0.6337 0.01644 0.01048 -0.0955 0.9333 0.0139
-9.250 -0.5994 0.01577 0.00974 -0.0973 0.9257 0.0150
-9.000 -0.5597 0.01502 0.00894 -0.1003 0.9174 0.0166
-8.750 -0.5141 0.01432 0.00819 -0.1045 0.9090 0.0196
-8.500 -0.4658 0.01363 0.00745 -0.1092 0.8964 0.0251
-8.250 -0.4249 0.01300 0.00677 -0.1124 0.8775 0.0345
-8.000 -0.3922 0.01248 0.00621 -0.1138 0.8543 0.0468
-7.750 -0.3645 0.01206 0.00574 -0.1142 0.8318 0.0602
-7.500 -0.3387 0.01171 0.00534 -0.1140 0.8104 0.0742
-7.250 -0.3136 0.01140 0.00499 -0.1136 0.7908 0.0881
-7.000 -0.2885 0.01113 0.00468 -0.1132 0.7724 0.1020
-6.750 -0.2633 0.01088 0.00440 -0.1128 0.7553 0.1175
-6.500 -0.2381 0.01061 0.00413 -0.1124 0.7392 0.1373
-6.250 -0.2126 0.01035 0.00389 -0.1121 0.7242 0.1604
-5.750 -0.1603 0.01004 0.00354 -0.1114 0.6959 0.1944
-5.500 -0.1339 0.00993 0.00339 -0.1111 0.6826 0.2068
-5.250 -0.1070 0.00985 0.00325 -0.1108 0.6697 0.2162
-5.000 -0.0800 0.00979 0.00313 -0.1106 0.6572 0.2255
-4.750 -0.0532 0.00973 0.00302 -0.1103 0.6449 0.2344
-4.500 -0.0259 0.00969 0.00291 -0.1100 0.6330 0.2427
-4.250 0.0013 0.00963 0.00282 -0.1098 0.6222 0.2507
-4.000 0.0285 0.00962 0.00274 -0.1095 0.6111 0.2588
-3.750 0.0558 0.00958 0.00266 -0.1093 0.5996 0.2663
-3.500 0.0832 0.00957 0.00259 -0.1091 0.5895 0.2743
-3.250 0.1106 0.00956 0.00253 -0.1089 0.5793 0.2810
-3.000 0.1381 0.00954 0.00248 -0.1087 0.5693 0.2887
-2.500 0.1930 0.00954 0.00240 -0.1083 0.5498 0.3029
-2.250 0.2203 0.00957 0.00236 -0.1081 0.5409 0.3096
-2.000 0.2480 0.00957 0.00234 -0.1080 0.5316 0.3166
-1.750 0.2753 0.00960 0.00232 -0.1077 0.5226 0.3235
-1.500 0.3031 0.00962 0.00231 -0.1076 0.5140 0.3298
-1.250 0.3302 0.00965 0.00231 -0.1074 0.5060 0.3370
-1.000 0.3580 0.00967 0.00231 -0.1072 0.4979 0.3438
-0.500 0.4129 0.00974 0.00234 -0.1068 0.4819 0.3577
-0.250 0.4403 0.00980 0.00236 -0.1066 0.4746 0.3642
0.000 0.4676 0.00984 0.00240 -0.1065 0.4679 0.3709
0.250 0.4951 0.00989 0.00243 -0.1063 0.4604 0.3786
0.500 0.5221 0.00996 0.00248 -0.1060 0.4533 0.3853
0.750 0.5497 0.01000 0.00252 -0.1059 0.4468 0.3926
1.000 0.5769 0.01007 0.00258 -0.1056 0.4403 0.3999
1.250 0.6039 0.01013 0.00265 -0.1054 0.4342 0.4077
1.500 0.6313 0.01019 0.00271 -0.1052 0.4277 0.4155
1.750 0.6580 0.01028 0.00279 -0.1049 0.4212 0.4229
2.000 0.6852 0.01034 0.00287 -0.1047 0.4160 0.4319
2.250 0.7124 0.01041 0.00295 -0.1045 0.4104 0.4402
2.500 0.7389 0.01051 0.00305 -0.1042 0.4046 0.4490
2.750 0.7658 0.01059 0.00315 -0.1040 0.3993 0.4583
3.000 0.7927 0.01066 0.00326 -0.1037 0.3936 0.4693
3.250 0.8190 0.01076 0.00338 -0.1034 0.3885 0.4794
3.500 0.8455 0.01087 0.00351 -0.1031 0.3839 0.4911
3.750 0.8722 0.01094 0.00363 -0.1028 0.3786 0.5042
4.000 0.8983 0.01104 0.00378 -0.1025 0.3732 0.5182
4.250 0.9240 0.01117 0.00393 -0.1020 0.3686 0.5329
4.500 0.9506 0.01124 0.00408 -0.1017 0.3642 0.5505
4.750 0.9766 0.01133 0.00424 -0.1014 0.3590 0.5703
5.000 1.0018 0.01146 0.00442 -0.1009 0.3539 0.5922
5.250 1.0274 0.01156 0.00460 -0.1004 0.3496 0.6179
5.500 1.0530 0.01163 0.00480 -0.1000 0.3445 0.6482
5.750 1.0775 0.01174 0.00500 -0.0993 0.3393 0.6823
6.000 1.1013 0.01185 0.00522 -0.0986 0.3346 0.7236
6.250 1.1252 0.01187 0.00544 -0.0977 0.3301 0.7769
6.500 1.1454 0.01188 0.00565 -0.0961 0.3250 0.8482
6.750 1.1703 0.01191 0.00582 -0.0954 0.3198 1.0000
7.000 1.1954 0.01210 0.00603 -0.0949 0.3151 1.0000
7.250 1.2196 0.01231 0.00625 -0.0943 0.3092 1.0000
7.500 1.2422 0.01258 0.00650 -0.0934 0.3034 1.0000
7.750 1.2663 0.01279 0.00673 -0.0928 0.2977 1.0000
8.000 1.2890 0.01304 0.00699 -0.0919 0.2914 1.0000
8.250 1.3103 0.01333 0.00726 -0.0908 0.2854 1.0000
8.500 1.3322 0.01356 0.00752 -0.0897 0.2790 1.0000
8.750 1.3515 0.01388 0.00783 -0.0883 0.2723 1.0000
9.000 1.3721 0.01416 0.00813 -0.0870 0.2662 1.0000
9.250 1.3902 0.01452 0.00848 -0.0854 0.2575 1.0000
9.500 1.4086 0.01488 0.00884 -0.0839 0.2475 1.0000
9.750 1.4249 0.01533 0.00926 -0.0821 0.2374 1.0000
10.000 1.4408 0.01581 0.00971 -0.0802 0.2264 1.0000
10.250 1.4557 0.01633 0.01021 -0.0783 0.2138 1.0000
10.500 1.4691 0.01693 0.01077 -0.0762 0.2018 1.0000
10.750 1.4799 0.01766 0.01145 -0.0738 0.1873 1.0000
11.000 1.4901 0.01843 0.01218 -0.0715 0.1746 1.0000
11.250 1.4986 0.01931 0.01302 -0.0690 0.1620 1.0000
11.500 1.5052 0.02033 0.01399 -0.0664 0.1492 1.0000
11.750 1.5095 0.02151 0.01513 -0.0638 0.1356 1.0000
12.000 1.5131 0.02282 0.01640 -0.0613 0.1237 1.0000
12.500 1.5181 0.02581 0.01936 -0.0569 0.1025 1.0000
12.750 1.5213 0.02742 0.02098 -0.0551 0.0951 1.0000
13.000 1.5213 0.02937 0.02293 -0.0533 0.0872 1.0000
13.250 1.5232 0.03130 0.02489 -0.0520 0.0804 1.0000
13.500 1.5233 0.03350 0.02712 -0.0507 0.0749 1.0000
13.750 1.5228 0.03585 0.02950 -0.0497 0.0692 1.0000
14.000 1.5217 0.03839 0.03209 -0.0489 0.0646 1.0000
14.250 1.5188 0.04123 0.03497 -0.0483 0.0595 1.0000
14.500 1.5161 0.04415 0.03795 -0.0479 0.0556 1.0000
14.750 1.5109 0.04746 0.04130 -0.0477 0.0513 1.0000
15.000 1.5066 0.05080 0.04471 -0.0477 0.0479 1.0000
15.250 1.5018 0.05430 0.04828 -0.0479 0.0448 1.0000
15.500 1.4945 0.05822 0.05226 -0.0483 0.0417 1.0000
15.750 1.4889 0.06206 0.05618 -0.0488 0.0389 1.0000
16.000 1.4807 0.06633 0.06050 -0.0495 0.0362 1.0000
16.250 1.4754 0.07033 0.06459 -0.0503 0.0342 1.0000
16.500 1.4670 0.07485 0.06918 -0.0514 0.0317 1.0000
16.750 1.4603 0.07924 0.07364 -0.0525 0.0297 1.0000
17.000 1.4531 0.08375 0.07823 -0.0538 0.0277 1.0000
17.250 1.4453 0.08848 0.08303 -0.0552 0.0259 1.0000
17.500 1.4386 0.09310 0.08773 -0.0566 0.0239 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 558 AIRFOIL (e558-il)