EPPLER 551 AIRFOIL (e551-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 551 AIRFOIL (e551-il) Reynolds number: 200,000 Max Cl/Cd: 70.03 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e551-il-200000-n5.txt Download as CSV file: xf-e551-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 551 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.250 -0.5404 0.11988 0.11572 -0.0516 1.0000 0.0091
-16.000 -0.5694 0.10881 0.10449 -0.0573 1.0000 0.0090
-15.750 -0.5909 0.10019 0.09573 -0.0618 1.0000 0.0088
-15.500 -0.6132 0.09202 0.08737 -0.0661 1.0000 0.0088
-15.250 -0.6319 0.08503 0.08019 -0.0697 1.0000 0.0088
-15.000 -0.6471 0.07911 0.07409 -0.0725 1.0000 0.0088
-14.750 -0.6569 0.07440 0.06924 -0.0745 1.0000 0.0088
-14.500 -0.6695 0.06946 0.06410 -0.0763 1.0000 0.0089
-14.250 -0.6764 0.06560 0.06009 -0.0775 1.0000 0.0090
-14.000 -0.6819 0.06209 0.05641 -0.0783 1.0000 0.0090
-13.750 -0.6851 0.05905 0.05324 -0.0788 1.0000 0.0091
-13.500 -0.6890 0.05603 0.05003 -0.0790 1.0000 0.0093
-13.250 -0.6902 0.05354 0.04746 -0.0789 1.0000 0.0094
-13.000 -0.6898 0.05133 0.04518 -0.0786 1.0000 0.0095
-12.750 -0.6892 0.04927 0.04306 -0.0782 1.0000 0.0096
-12.500 -0.6888 0.04736 0.04115 -0.0779 1.0000 0.0100
-12.250 -0.6877 0.04554 0.03926 -0.0771 1.0000 0.0103
-12.000 -0.6866 0.04387 0.03752 -0.0760 1.0000 0.0105
-11.750 -0.6862 0.04233 0.03593 -0.0747 1.0000 0.0109
-11.250 -0.6590 0.03874 0.03211 -0.0764 0.9908 0.0118
-11.000 -0.6421 0.03681 0.03017 -0.0783 0.9769 0.0123
-10.750 -0.6230 0.03500 0.02831 -0.0803 0.9637 0.0128
-10.500 -0.6002 0.03318 0.02639 -0.0829 0.9530 0.0135
-10.250 -0.5739 0.03136 0.02445 -0.0862 0.9430 0.0141
-10.000 -0.5476 0.02942 0.02242 -0.0899 0.9317 0.0150
-9.750 -0.5187 0.02766 0.02057 -0.0938 0.9198 0.0161
-9.500 -0.4902 0.02601 0.01877 -0.0974 0.9068 0.0178
-9.000 -0.4486 0.02305 0.01563 -0.1013 0.8765 0.0226
-8.750 -0.4374 0.02186 0.01430 -0.1010 0.8610 0.0250
-8.500 -0.4319 0.02086 0.01321 -0.0994 0.8466 0.0278
-8.250 -0.4205 0.02003 0.01231 -0.0981 0.8345 0.0320
-8.000 -0.4070 0.01926 0.01147 -0.0969 0.8239 0.0394
-7.750 -0.3931 0.01855 0.01073 -0.0956 0.8132 0.0492
-7.500 -0.3781 0.01780 0.01001 -0.0945 0.8038 0.0636
-7.250 -0.3620 0.01704 0.00929 -0.0935 0.7945 0.0857
-7.000 -0.3455 0.01626 0.00861 -0.0926 0.7857 0.1171
-6.750 -0.3298 0.01529 0.00784 -0.0917 0.7774 0.1684
-6.500 -0.3161 0.01403 0.00698 -0.0908 0.7691 0.2582
-6.250 -0.2976 0.01336 0.00679 -0.0900 0.7612 0.3704
-6.000 -0.2713 0.01336 0.00675 -0.0898 0.7540 0.4071
-5.750 -0.2443 0.01340 0.00664 -0.0897 0.7466 0.4280
-5.500 -0.2169 0.01348 0.00658 -0.0896 0.7399 0.4429
-5.250 -0.1899 0.01357 0.00661 -0.0894 0.7323 0.4541
-5.000 -0.1623 0.01365 0.00654 -0.0893 0.7259 0.4642
-4.750 -0.1353 0.01382 0.00660 -0.0891 0.7189 0.4775
-4.500 -0.1079 0.01407 0.00682 -0.0888 0.7124 0.4884
-4.250 -0.0803 0.01406 0.00664 -0.0888 0.7062 0.4945
-4.000 -0.0529 0.01399 0.00651 -0.0888 0.6992 0.4968
-3.750 -0.0250 0.01393 0.00634 -0.0888 0.6933 0.4992
-3.500 0.0026 0.01386 0.00619 -0.0889 0.6869 0.5017
-3.250 0.0303 0.01380 0.00602 -0.0890 0.6809 0.5043
-2.750 0.0860 0.01368 0.00564 -0.0893 0.6691 0.5099
-2.500 0.1137 0.01362 0.00552 -0.0893 0.6631 0.5117
-2.250 0.1416 0.01359 0.00542 -0.0894 0.6580 0.5136
-2.000 0.1690 0.01356 0.00537 -0.0895 0.6521 0.5159
-1.750 0.1969 0.01355 0.00529 -0.0895 0.6467 0.5185
-1.500 0.2249 0.01354 0.00519 -0.0897 0.6419 0.5211
-1.250 0.2524 0.01352 0.00513 -0.0898 0.6358 0.5236
-1.000 0.2803 0.01352 0.00504 -0.0899 0.6305 0.5262
-0.500 0.3356 0.01352 0.00500 -0.0901 0.6209 0.5305
-0.250 0.3631 0.01354 0.00501 -0.0901 0.6157 0.5331
0.000 0.3910 0.01357 0.00499 -0.0902 0.6111 0.5360
0.250 0.4185 0.01360 0.00500 -0.0903 0.6062 0.5388
0.500 0.4460 0.01363 0.00502 -0.0904 0.6013 0.5417
0.750 0.4738 0.01367 0.00502 -0.0905 0.5969 0.5444
1.000 0.5015 0.01371 0.00505 -0.0906 0.5928 0.5467
1.250 0.5283 0.01376 0.00515 -0.0906 0.5877 0.5494
1.500 0.5555 0.01381 0.00522 -0.0906 0.5831 0.5526
1.750 0.5833 0.01389 0.00527 -0.0907 0.5791 0.5562
2.000 0.6107 0.01397 0.00534 -0.0908 0.5748 0.5599
2.250 0.6373 0.01403 0.00547 -0.0907 0.5699 0.5628
2.500 0.6643 0.01410 0.00558 -0.0907 0.5656 0.5657
2.750 0.6920 0.01419 0.00566 -0.0908 0.5618 0.5691
3.000 0.7184 0.01429 0.00581 -0.0907 0.5571 0.5729
3.250 0.7450 0.01439 0.00595 -0.0906 0.5524 0.5770
3.500 0.7717 0.01447 0.00608 -0.0906 0.5481 0.5807
3.750 0.7989 0.01458 0.00622 -0.0906 0.5441 0.5848
4.000 0.8242 0.01469 0.00643 -0.0903 0.5388 0.5895
4.250 0.8503 0.01481 0.00658 -0.0901 0.5338 0.5945
4.500 0.8770 0.01491 0.00672 -0.0900 0.5293 0.5985
4.750 0.9017 0.01503 0.00695 -0.0896 0.5237 0.6032
5.000 0.9268 0.01514 0.00714 -0.0893 0.5179 0.6085
5.250 0.9530 0.01527 0.00727 -0.0891 0.5129 0.6138
5.500 0.9764 0.01539 0.00754 -0.0884 0.5065 0.6193
5.750 1.0008 0.01552 0.00774 -0.0880 0.5004 0.6260
6.000 1.0255 0.01566 0.00794 -0.0875 0.4948 0.6324
6.250 1.0481 0.01580 0.00822 -0.0867 0.4878 0.6393
6.500 1.0720 0.01595 0.00841 -0.0861 0.4813 0.6468
6.750 1.0935 0.01610 0.00871 -0.0852 0.4736 0.6540
7.000 1.1155 0.01627 0.00892 -0.0843 0.4659 0.6625
7.250 1.1359 0.01643 0.00924 -0.0831 0.4576 0.6710
7.500 1.1562 0.01662 0.00948 -0.0819 0.4491 0.6810
7.750 1.1744 0.01681 0.00982 -0.0803 0.4394 0.6919
8.000 1.1919 0.01702 0.01011 -0.0786 0.4296 0.7035
8.250 1.2068 0.01725 0.01044 -0.0765 0.4184 0.7164
8.500 1.2193 0.01749 0.01080 -0.0739 0.4066 0.7306
9.000 1.2406 0.01820 0.01167 -0.0684 0.3792 0.7651
9.250 1.2483 0.01867 0.01221 -0.0653 0.3630 0.7886
9.500 1.2530 0.01922 0.01284 -0.0618 0.3455 0.8210
9.750 1.2538 0.01980 0.01354 -0.0577 0.3275 0.8903
10.250 1.2585 0.02189 0.01553 -0.0521 0.2879 1.0000
10.500 1.2591 0.02326 0.01682 -0.0495 0.2685 1.0000
10.750 1.2590 0.02478 0.01829 -0.0471 0.2504 1.0000
11.000 1.2587 0.02645 0.01990 -0.0448 0.2334 1.0000
11.250 1.2582 0.02824 0.02164 -0.0428 0.2172 1.0000
11.500 1.2578 0.03013 0.02350 -0.0411 0.2021 1.0000
11.750 1.2574 0.03213 0.02546 -0.0395 0.1875 1.0000
12.000 1.2574 0.03419 0.02751 -0.0382 0.1739 1.0000
12.250 1.2573 0.03635 0.02965 -0.0370 0.1608 1.0000
12.500 1.2572 0.03860 0.03190 -0.0359 0.1489 1.0000
12.750 1.2571 0.04094 0.03423 -0.0350 0.1382 1.0000
13.000 1.2559 0.04346 0.03674 -0.0343 0.1280 1.0000
13.250 1.2574 0.04583 0.03912 -0.0337 0.1180 1.0000
13.500 1.2581 0.04834 0.04166 -0.0333 0.1091 1.0000
13.750 1.2574 0.05109 0.04440 -0.0329 0.1008 1.0000
14.000 1.2585 0.05371 0.04705 -0.0327 0.0925 1.0000
14.250 1.2591 0.05646 0.04983 -0.0326 0.0853 1.0000
14.500 1.2583 0.05946 0.05284 -0.0326 0.0785 1.0000
14.750 1.2601 0.06221 0.05565 -0.0327 0.0721 1.0000
15.000 1.2588 0.06539 0.05884 -0.0329 0.0665 1.0000
15.250 1.2603 0.06831 0.06182 -0.0332 0.0611 1.0000
15.500 1.2594 0.07159 0.06514 -0.0336 0.0566 1.0000
15.750 1.2601 0.07473 0.06834 -0.0341 0.0520 1.0000
16.000 1.2584 0.07824 0.07190 -0.0348 0.0484 1.0000
16.250 1.2592 0.08149 0.07523 -0.0355 0.0447 1.0000
16.500 1.2566 0.08528 0.07907 -0.0364 0.0418 1.0000
16.750 1.2571 0.08868 0.08257 -0.0372 0.0387 1.0000
17.000 1.2552 0.09247 0.08642 -0.0383 0.0363 1.0000
17.250 1.2530 0.09637 0.09040 -0.0395 0.0341 1.0000
17.500 1.2520 0.10016 0.09430 -0.0407 0.0318 1.0000
17.750 1.2485 0.10438 0.09857 -0.0422 0.0300 1.0000
18.000 1.2466 0.10839 0.10268 -0.0436 0.0283 1.0000
18.250 1.2446 0.11247 0.10687 -0.0452 0.0266 1.0000
18.500 1.2405 0.11695 0.11141 -0.0471 0.0253 1.0000
18.750 1.2376 0.12126 0.11582 -0.0490 0.0239 1.0000
19.000 1.2355 0.12548 0.12017 -0.0509 0.0225 1.0000
19.250 1.2317 0.13003 0.12481 -0.0531 0.0214 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 551 AIRFOIL (e551-il)