EPPLER 502 AIRFOIL (e502-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 502 AIRFOIL (e502-il) Reynolds number: 500,000 Max Cl/Cd: 88.01 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e502-il-500000.txt Download as CSV file: xf-e502-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 502 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.750 -0.3318 0.12448 0.12199 -0.0421 0.8436 0.0174
-13.500 -0.3325 0.11988 0.11736 -0.0445 0.8395 0.0174
-13.250 -0.4331 0.12163 0.11945 -0.0424 1.0000 0.0173
-11.750 -0.5756 0.04026 0.03686 -0.0763 0.8312 0.0088
-11.500 -0.5847 0.03748 0.03387 -0.0753 0.8251 0.0086
-11.250 -0.5933 0.03491 0.03113 -0.0739 0.8190 0.0085
-11.000 -0.5987 0.03296 0.02901 -0.0722 0.8135 0.0084
-10.750 -0.6641 0.03833 0.03393 -0.0712 0.8185 0.0081
-10.250 -0.6853 0.03242 0.02741 -0.0635 0.8060 0.0077
-10.000 -0.6879 0.02922 0.02377 -0.0599 0.8009 0.0077
-9.750 -0.6785 0.02705 0.02130 -0.0578 0.7968 0.0074
-9.500 -0.6627 0.02479 0.01872 -0.0565 0.7928 0.0075
-9.250 -0.6429 0.02304 0.01672 -0.0557 0.7891 0.0076
-9.000 -0.6222 0.02164 0.01511 -0.0548 0.7858 0.0079
-8.750 -0.6022 0.02065 0.01396 -0.0539 0.7827 0.0081
-8.500 -0.5830 0.01961 0.01281 -0.0529 0.7792 0.0081
-8.250 -0.5679 0.01833 0.01138 -0.0512 0.7756 0.0089
-8.000 -0.5526 0.01751 0.01048 -0.0495 0.7724 0.0094
-7.750 -0.5346 0.01701 0.00993 -0.0482 0.7696 0.0104
-7.500 -0.5169 0.01647 0.00933 -0.0468 0.7668 0.0118
-7.250 -0.5031 0.01571 0.00855 -0.0447 0.7638 0.0142
-7.000 -0.4829 0.01541 0.00822 -0.0435 0.7609 0.0175
-6.750 -0.4687 0.01485 0.00765 -0.0414 0.7582 0.0225
-6.500 -0.4518 0.01451 0.00728 -0.0398 0.7556 0.0281
-6.250 -0.4312 0.01435 0.00710 -0.0387 0.7532 0.0325
-6.000 -0.4152 0.01391 0.00666 -0.0368 0.7507 0.0378
-5.750 -0.3941 0.01366 0.00640 -0.0359 0.7482 0.0430
-5.500 -0.3761 0.01325 0.00597 -0.0343 0.7458 0.0498
-5.250 -0.3564 0.01292 0.00562 -0.0331 0.7435 0.0591
-5.000 -0.3391 0.01247 0.00524 -0.0314 0.7413 0.0807
-4.750 -0.3252 0.01185 0.00490 -0.0292 0.7390 0.1388
-4.500 -0.3154 0.01102 0.00454 -0.0265 0.7366 0.2403
-4.250 -0.3093 0.01002 0.00412 -0.0230 0.7342 0.3739
-4.000 -0.3110 0.00869 0.00362 -0.0179 0.7318 0.5632
-3.750 -0.2936 0.00855 0.00385 -0.0157 0.7297 0.6779
-3.500 -0.2671 0.00870 0.00392 -0.0153 0.7277 0.7016
-3.250 -0.2401 0.00889 0.00404 -0.0151 0.7259 0.7164
-3.000 -0.2133 0.00907 0.00420 -0.0148 0.7239 0.7295
-2.750 -0.1865 0.00925 0.00433 -0.0146 0.7218 0.7406
-2.500 -0.1582 0.00949 0.00457 -0.0145 0.7196 0.7501
-2.250 -0.1301 0.00980 0.00489 -0.0142 0.7175 0.7613
-2.000 -0.1022 0.01013 0.00522 -0.0139 0.7156 0.7722
-1.750 -0.0737 0.01044 0.00549 -0.0138 0.7139 0.7809
-1.500 -0.0453 0.01062 0.00564 -0.0138 0.7123 0.7861
-1.250 -0.0170 0.01069 0.00571 -0.0139 0.7102 0.7891
-1.000 0.0107 0.01066 0.00566 -0.0142 0.7079 0.7909
-0.750 0.0385 0.01062 0.00559 -0.0145 0.7057 0.7927
-0.500 0.0664 0.01057 0.00550 -0.0149 0.7036 0.7943
-0.250 0.0945 0.01053 0.00540 -0.0153 0.7016 0.7959
0.000 0.1228 0.01050 0.00531 -0.0157 0.6997 0.7972
0.250 0.1515 0.01048 0.00525 -0.0162 0.6979 0.7983
0.500 0.1791 0.01045 0.00524 -0.0164 0.6955 0.7993
0.750 0.2069 0.01041 0.00522 -0.0167 0.6928 0.8003
1.000 0.2351 0.01038 0.00519 -0.0170 0.6901 0.8013
1.250 0.2635 0.01034 0.00514 -0.0174 0.6875 0.8025
1.500 0.2925 0.01032 0.00509 -0.0179 0.6851 0.8036
1.750 0.3209 0.01033 0.00508 -0.0183 0.6824 0.8047
2.000 0.3481 0.01028 0.00508 -0.0185 0.6788 0.8059
2.250 0.3760 0.01022 0.00503 -0.0188 0.6752 0.8071
2.500 0.4049 0.01016 0.00495 -0.0192 0.6719 0.8083
2.750 0.4339 0.01016 0.00491 -0.0197 0.6689 0.8097
3.000 0.4605 0.01011 0.00492 -0.0198 0.6644 0.8110
3.250 0.4885 0.01002 0.00486 -0.0200 0.6601 0.8120
3.500 0.5175 0.00994 0.00477 -0.0205 0.6564 0.8129
3.750 0.5447 0.00989 0.00478 -0.0206 0.6519 0.8139
4.000 0.5719 0.00982 0.00477 -0.0207 0.6467 0.8150
4.250 0.6006 0.00974 0.00469 -0.0210 0.6420 0.8161
4.500 0.6269 0.00968 0.00470 -0.0209 0.6358 0.8174
4.750 0.6543 0.00958 0.00464 -0.0210 0.6293 0.8188
5.000 0.6808 0.00953 0.00466 -0.0209 0.6222 0.8205
5.250 0.7075 0.00946 0.00463 -0.0209 0.6147 0.8220
5.500 0.7334 0.00941 0.00464 -0.0208 0.6053 0.8235
5.750 0.7593 0.00937 0.00464 -0.0206 0.5946 0.8250
6.000 0.7841 0.00933 0.00465 -0.0202 0.5794 0.8264
6.250 0.8076 0.00933 0.00467 -0.0195 0.5599 0.8279
6.500 0.8291 0.00942 0.00474 -0.0185 0.5323 0.8295
6.750 0.8451 0.00972 0.00489 -0.0164 0.4886 0.8314
7.000 0.8549 0.01026 0.00521 -0.0134 0.4355 0.8337
7.250 0.8624 0.01089 0.00564 -0.0101 0.3885 0.8363
7.500 0.8669 0.01157 0.00612 -0.0062 0.3432 0.8393
7.750 0.8659 0.01218 0.00657 -0.0014 0.3027 0.8423
8.000 0.8645 0.01291 0.00715 0.0034 0.2644 0.8456
8.250 0.8650 0.01371 0.00782 0.0075 0.2312 0.8493
8.500 0.8667 0.01458 0.00855 0.0110 0.1995 0.8533
8.750 0.8693 0.01548 0.00934 0.0142 0.1704 0.8573
9.000 0.8724 0.01645 0.01021 0.0172 0.1429 0.8618
9.250 0.8768 0.01746 0.01112 0.0198 0.1188 0.8670
9.500 0.8820 0.01845 0.01204 0.0221 0.0962 0.8722
9.750 0.8877 0.01945 0.01300 0.0243 0.0772 0.8788
10.000 0.8942 0.02046 0.01397 0.0264 0.0609 0.8864
10.250 0.9014 0.02144 0.01495 0.0283 0.0490 0.8964
10.500 0.9089 0.02243 0.01596 0.0301 0.0381 0.9101
10.750 0.9233 0.02346 0.01707 0.0305 0.0294 0.9324
11.000 0.9526 0.02507 0.01871 0.0268 0.0202 0.9590
11.250 0.9724 0.02655 0.02021 0.0253 0.0146 1.0000
11.500 0.9759 0.02819 0.02186 0.0269 0.0120 1.0000
11.750 0.9858 0.02940 0.02313 0.0278 0.0099 1.0000
12.000 0.9927 0.03083 0.02460 0.0289 0.0093 1.0000
12.250 0.9903 0.03301 0.02686 0.0307 0.0082 1.0000
12.500 0.9987 0.03441 0.02834 0.0316 0.0075 1.0000
12.750 1.0030 0.03617 0.03019 0.0327 0.0074 1.0000
13.000 1.0116 0.03762 0.03169 0.0334 0.0066 1.0000
13.250 1.0159 0.03948 0.03360 0.0342 0.0062 1.0000
13.500 1.0186 0.04152 0.03574 0.0351 0.0062 1.0000
13.750 1.0167 0.04406 0.03835 0.0360 0.0058 1.0000
14.000 1.0101 0.04716 0.04159 0.0372 0.0053 1.0000
14.250 1.0130 0.04942 0.04395 0.0378 0.0053 1.0000
14.500 1.0138 0.05190 0.04651 0.0383 0.0055 1.0000
14.750 1.0179 0.05421 0.04896 0.0385 0.0053 1.0000
15.000 1.0194 0.05686 0.05176 0.0387 0.0051 1.0000
15.250 1.0194 0.05971 0.05475 0.0389 0.0051 1.0000
15.500 1.0184 0.06280 0.05798 0.0388 0.0051 1.0000
15.750 1.0166 0.06609 0.06141 0.0386 0.0051 1.0000
16.000 1.0132 0.06968 0.06516 0.0380 0.0050 1.0000
16.250 1.0070 0.07378 0.06943 0.0373 0.0050 1.0000
16.500 0.9986 0.07839 0.07422 0.0362 0.0050 1.0000
16.750 0.9908 0.08306 0.07905 0.0347 0.0050 1.0000
17.000 0.9775 0.08879 0.08498 0.0327 0.0050 1.0000
17.250 0.9659 0.09448 0.09084 0.0304 0.0051 1.0000
17.500 0.9536 0.10055 0.09707 0.0276 0.0051 1.0000
17.750 0.9405 0.10696 0.10362 0.0244 0.0052 1.0000
18.000 0.9204 0.11511 0.11197 0.0202 0.0052 1.0000
18.250 0.9080 0.12194 0.11893 0.0164 0.0052 1.0000
18.500 0.8874 0.13095 0.12811 0.0113 0.0053 1.0000
18.750 0.8688 0.13974 0.13703 0.0063 0.0054 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 502 AIRFOIL (e502-il)