EPPLER 502 AIRFOIL (e502-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 502 AIRFOIL (e502-il) Reynolds number: 1,000,000 Max Cl/Cd: 90.6 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e502-il-1000000-n5.txt Download as CSV file: xf-e502-il-1000000-n5.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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.000 -0.7685 0.08911 0.08710 -0.0471 0.9997 0.0019
-15.750 -0.7994 0.07441 0.07219 -0.0576 0.9624 0.0019
-15.500 -0.8213 0.06038 0.05781 -0.0722 0.9411 0.0018
-15.250 -0.8276 0.05264 0.04955 -0.0804 0.8717 0.0018
-15.000 -0.8455 0.04852 0.04513 -0.0803 0.8373 0.0018
-14.750 -0.8540 0.04553 0.04195 -0.0799 0.8196 0.0018
-14.500 -0.8731 0.04160 0.03776 -0.0789 0.8071 0.0018
-14.250 -0.8789 0.03915 0.03513 -0.0777 0.7969 0.0018
-14.000 -0.8791 0.03723 0.03307 -0.0766 0.7881 0.0018
-13.750 -0.8857 0.03479 0.03043 -0.0748 0.7808 0.0018
-13.500 -0.8856 0.03296 0.02844 -0.0732 0.7739 0.0018
-13.250 -0.8849 0.03122 0.02654 -0.0714 0.7678 0.0018
-13.000 -0.8832 0.02953 0.02468 -0.0695 0.7619 0.0018
-12.750 -0.8792 0.02805 0.02304 -0.0676 0.7565 0.0019
-12.500 -0.8756 0.02650 0.02131 -0.0656 0.7518 0.0019
-12.250 -0.8660 0.02544 0.02013 -0.0640 0.7473 0.0019
-12.000 -0.8553 0.02448 0.01905 -0.0625 0.7431 0.0019
-11.750 -0.8465 0.02330 0.01772 -0.0606 0.7395 0.0019
-11.500 -0.8345 0.02237 0.01670 -0.0591 0.7361 0.0019
-11.250 -0.8220 0.02147 0.01568 -0.0575 0.7326 0.0019
-11.000 -0.8082 0.02068 0.01479 -0.0560 0.7292 0.0020
-10.750 -0.7936 0.01996 0.01396 -0.0546 0.7260 0.0020
-10.500 -0.7791 0.01917 0.01308 -0.0531 0.7231 0.0020
-10.250 -0.7635 0.01848 0.01231 -0.0517 0.7200 0.0020
-10.000 -0.7478 0.01779 0.01152 -0.0502 0.7173 0.0021
-9.750 -0.7299 0.01732 0.01099 -0.0491 0.7146 0.0020
-9.250 -0.6976 0.01600 0.00951 -0.0461 0.7101 0.0022
-9.000 -0.6812 0.01537 0.00881 -0.0445 0.7077 0.0025
-8.750 -0.6627 0.01495 0.00835 -0.0433 0.7052 0.0026
-8.500 -0.6440 0.01455 0.00789 -0.0420 0.7027 0.0025
-8.250 -0.6260 0.01413 0.00743 -0.0406 0.7003 0.0029
-8.000 -0.6075 0.01377 0.00700 -0.0393 0.6983 0.0029
-7.750 -0.5890 0.01342 0.00661 -0.0379 0.6965 0.0032
-7.500 -0.5705 0.01310 0.00626 -0.0364 0.6946 0.0035
-7.250 -0.5531 0.01277 0.00590 -0.0347 0.6926 0.0043
-7.000 -0.5344 0.01248 0.00560 -0.0332 0.6904 0.0051
-6.750 -0.5137 0.01223 0.00531 -0.0322 0.6882 0.0058
-6.500 -0.4932 0.01193 0.00501 -0.0311 0.6863 0.0084
-6.250 -0.4719 0.01166 0.00473 -0.0301 0.6845 0.0113
-6.000 -0.4499 0.01142 0.00450 -0.0293 0.6829 0.0142
-5.750 -0.4268 0.01121 0.00429 -0.0287 0.6813 0.0170
-5.500 -0.4037 0.01098 0.00407 -0.0280 0.6795 0.0200
-5.250 -0.3807 0.01076 0.00387 -0.0274 0.6776 0.0255
-5.000 -0.3578 0.01052 0.00366 -0.0267 0.6757 0.0319
-4.750 -0.3345 0.01030 0.00347 -0.0261 0.6739 0.0401
-4.500 -0.3108 0.01010 0.00328 -0.0256 0.6722 0.0491
-4.250 -0.2872 0.00989 0.00310 -0.0250 0.6706 0.0613
-4.000 -0.2640 0.00966 0.00293 -0.0244 0.6690 0.0801
-3.500 -0.2192 0.00903 0.00256 -0.0229 0.6656 0.1506
-3.250 -0.1971 0.00869 0.00238 -0.0221 0.6639 0.1959
-2.750 -0.1551 0.00788 0.00201 -0.0202 0.6605 0.3227
-2.500 -0.1413 0.00703 0.00168 -0.0180 0.6587 0.4638
-2.250 -0.1300 0.00610 0.00143 -0.0152 0.6569 0.6438
-2.000 -0.1034 0.00609 0.00146 -0.0150 0.6551 0.6816
-1.750 -0.0756 0.00611 0.00151 -0.0151 0.6537 0.7008
-1.500 -0.0472 0.00616 0.00154 -0.0154 0.6521 0.7122
-1.250 -0.0188 0.00619 0.00159 -0.0156 0.6503 0.7199
-1.000 0.0098 0.00624 0.00160 -0.0159 0.6484 0.7257
-0.750 0.0386 0.00624 0.00160 -0.0162 0.6466 0.7290
-0.500 0.0675 0.00625 0.00159 -0.0166 0.6448 0.7306
-0.250 0.0965 0.00626 0.00157 -0.0171 0.6429 0.7315
0.000 0.1253 0.00627 0.00155 -0.0175 0.6410 0.7324
0.250 0.1544 0.00627 0.00154 -0.0179 0.6392 0.7332
0.500 0.1835 0.00627 0.00154 -0.0184 0.6367 0.7341
0.750 0.2124 0.00627 0.00153 -0.0188 0.6336 0.7352
1.000 0.2412 0.00628 0.00152 -0.0192 0.6302 0.7362
1.250 0.2697 0.00631 0.00152 -0.0196 0.6269 0.7370
1.500 0.2987 0.00630 0.00153 -0.0200 0.6233 0.7378
1.750 0.3275 0.00631 0.00154 -0.0205 0.6189 0.7386
2.000 0.3560 0.00633 0.00154 -0.0208 0.6146 0.7393
2.250 0.3846 0.00635 0.00156 -0.0212 0.6104 0.7401
2.500 0.4131 0.00635 0.00159 -0.0216 0.6050 0.7409
2.750 0.4411 0.00638 0.00161 -0.0218 0.5993 0.7418
3.000 0.4696 0.00640 0.00165 -0.0222 0.5934 0.7427
3.250 0.4975 0.00643 0.00169 -0.0224 0.5864 0.7437
3.500 0.5255 0.00647 0.00174 -0.0227 0.5780 0.7448
3.750 0.5528 0.00653 0.00179 -0.0228 0.5671 0.7458
4.000 0.5796 0.00662 0.00186 -0.0228 0.5526 0.7468
4.250 0.6053 0.00675 0.00195 -0.0227 0.5322 0.7478
4.500 0.6297 0.00695 0.00206 -0.0223 0.5056 0.7488
4.750 0.6520 0.00725 0.00224 -0.0215 0.4717 0.7499
5.000 0.6718 0.00768 0.00249 -0.0204 0.4279 0.7510
5.250 0.6907 0.00813 0.00278 -0.0190 0.3851 0.7521
5.500 0.7088 0.00860 0.00307 -0.0176 0.3435 0.7532
6.000 0.7429 0.00953 0.00370 -0.0144 0.2661 0.7556
6.250 0.7578 0.01003 0.00404 -0.0124 0.2295 0.7569
6.500 0.7730 0.01047 0.00436 -0.0104 0.1996 0.7583
6.750 0.7875 0.01089 0.00469 -0.0083 0.1741 0.7600
7.000 0.7998 0.01126 0.00499 -0.0058 0.1533 0.7617
7.250 0.8082 0.01171 0.00535 -0.0025 0.1316 0.7635
7.500 0.8171 0.01224 0.00580 0.0005 0.1101 0.7653
7.750 0.8267 0.01279 0.00628 0.0032 0.0925 0.7669
8.000 0.8377 0.01333 0.00676 0.0056 0.0779 0.7685
8.250 0.8473 0.01394 0.00732 0.0080 0.0627 0.7704
8.500 0.8565 0.01461 0.00794 0.0104 0.0487 0.7724
8.750 0.8639 0.01542 0.00867 0.0129 0.0338 0.7745
9.000 0.8773 0.01602 0.00928 0.0145 0.0284 0.7767
9.250 0.8888 0.01674 0.00997 0.0162 0.0215 0.7790
9.500 0.9011 0.01744 0.01068 0.0177 0.0170 0.7812
10.000 0.9268 0.01885 0.01211 0.0205 0.0107 0.7859
10.250 0.9411 0.01950 0.01280 0.0217 0.0100 0.7888
10.500 0.9528 0.02032 0.01363 0.0231 0.0073 0.7921
10.750 0.9648 0.02113 0.01447 0.0244 0.0056 0.7952
11.000 0.9781 0.02190 0.01528 0.0255 0.0050 0.7982
11.250 0.9899 0.02275 0.01617 0.0267 0.0040 0.8017
11.500 1.0021 0.02358 0.01707 0.0279 0.0036 0.8056
11.750 1.0138 0.02448 0.01803 0.0290 0.0030 0.8096
12.000 1.0261 0.02534 0.01895 0.0300 0.0029 0.8137
12.250 1.0369 0.02631 0.01998 0.0312 0.0025 0.8187
12.500 1.0455 0.02744 0.02117 0.0325 0.0021 0.8243
12.750 1.0568 0.02840 0.02221 0.0335 0.0022 0.8302
13.000 1.0666 0.02948 0.02337 0.0345 0.0018 0.8377
13.250 1.0761 0.03057 0.02455 0.0356 0.0017 0.8461
13.500 1.0848 0.03173 0.02582 0.0366 0.0018 0.8570
13.750 1.0922 0.03295 0.02719 0.0379 0.0017 0.8726
14.000 1.0964 0.03432 0.02875 0.0394 0.0013 0.9062
14.250 1.1288 0.03618 0.03083 0.0344 0.0013 0.9552
14.500 1.1536 0.03772 0.03246 0.0316 0.0015 1.0000
14.750 1.1586 0.03952 0.03433 0.0323 0.0013 1.0000
15.000 1.1613 0.04157 0.03645 0.0331 0.0012 1.0000
15.250 1.1673 0.04338 0.03833 0.0335 0.0012 1.0000
15.500 1.1678 0.04575 0.04081 0.0340 0.0011 1.0000
15.750 1.1730 0.04775 0.04288 0.0343 0.0011 1.0000
16.000 1.1723 0.05040 0.04563 0.0346 0.0011 1.0000
16.250 1.1731 0.05297 0.04829 0.0347 0.0011 1.0000
16.500 1.1612 0.05708 0.05255 0.0347 0.0010 1.0000
16.750 1.1649 0.05953 0.05509 0.0344 0.0010 1.0000
17.000 1.1548 0.06369 0.05940 0.0339 0.0010 1.0000
17.250 1.1490 0.06753 0.06334 0.0332 0.0010 1.0000
17.500 1.1505 0.07049 0.06639 0.0324 0.0010 1.0000
17.750 1.1505 0.07376 0.06973 0.0315 0.0010 1.0000
18.000 1.1344 0.07938 0.07551 0.0298 0.0010 1.0000
18.250 1.1283 0.08376 0.08000 0.0283 0.0010 1.0000
18.500 1.1187 0.08884 0.08519 0.0263 0.0010 1.0000
18.750 1.1033 0.09502 0.09151 0.0238 0.0010 1.0000
19.000 1.0938 0.10038 0.09699 0.0214 0.0010 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 502 AIRFOIL (e502-il)