EPPLER 341 AIRFOIL (e341-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 341 AIRFOIL (e341-il) Reynolds number: 500,000 Max Cl/Cd: 80.67 at α=10.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e341-il-500000-n5.txt Download as CSV file: xf-e341-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 341 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5466 0.07160 0.06850 -0.0172 0.6715 0.0078
-8.750 -0.5627 0.06675 0.06351 -0.0170 0.6608 0.0077
-8.500 -0.5758 0.06218 0.05876 -0.0155 0.6502 0.0076
-8.250 -0.5828 0.05735 0.05373 -0.0139 0.6401 0.0075
-7.250 -0.6086 0.02772 0.02189 0.0022 0.6118 0.0066
-7.000 -0.5924 0.02535 0.01914 0.0041 0.6005 0.0065
-6.750 -0.5730 0.02338 0.01681 0.0055 0.5892 0.0065
-6.500 -0.5513 0.02180 0.01491 0.0067 0.5779 0.0065
-6.250 -0.5279 0.02050 0.01334 0.0075 0.5663 0.0065
-6.000 -0.5035 0.01936 0.01198 0.0082 0.5557 0.0065
-5.750 -0.4784 0.01842 0.01084 0.0088 0.5462 0.0066
-5.500 -0.4530 0.01758 0.00984 0.0094 0.5364 0.0066
-5.250 -0.4276 0.01676 0.00888 0.0099 0.5277 0.0066
-5.000 -0.4027 0.01599 0.00799 0.0105 0.5183 0.0067
-4.750 -0.3780 0.01532 0.00724 0.0111 0.5094 0.0069
-4.500 -0.3531 0.01482 0.00665 0.0117 0.5003 0.0070
-4.250 -0.3279 0.01438 0.00615 0.0123 0.4922 0.0072
-4.000 -0.3028 0.01400 0.00570 0.0129 0.4844 0.0075
-3.750 -0.2775 0.01364 0.00529 0.0134 0.4774 0.0078
-3.500 -0.2522 0.01332 0.00490 0.0140 0.4699 0.0081
-3.250 -0.2267 0.01304 0.00455 0.0145 0.4630 0.0087
-3.000 -0.2013 0.01277 0.00422 0.0150 0.4556 0.0092
-2.750 -0.1758 0.01254 0.00394 0.0155 0.4491 0.0102
-2.500 -0.1499 0.01234 0.00370 0.0160 0.4429 0.0114
-2.250 -0.1243 0.01214 0.00347 0.0165 0.4369 0.0140
-2.000 -0.0987 0.01193 0.00328 0.0170 0.4317 0.0213
-1.750 -0.0733 0.01170 0.00311 0.0174 0.4261 0.0383
-1.500 -0.0483 0.01149 0.00297 0.0180 0.4204 0.0659
-1.250 -0.0243 0.01118 0.00285 0.0186 0.4153 0.1213
-1.000 -0.0152 0.00979 0.00257 0.0216 0.4110 0.4237
-0.750 -0.0185 0.00852 0.00304 0.0289 0.4075 0.8527
-0.500 0.0030 0.00878 0.00322 0.0306 0.4031 0.8788
-0.250 0.0272 0.00903 0.00340 0.0317 0.3992 0.8916
0.000 0.0592 0.00955 0.00388 0.0316 0.3945 0.9006
0.250 0.0847 0.00985 0.00410 0.0326 0.3899 0.9092
0.500 0.1216 0.01034 0.00451 0.0314 0.3853 0.9141
0.750 0.1503 0.01049 0.00461 0.0314 0.3814 0.9181
1.000 0.1724 0.01046 0.00452 0.0325 0.3778 0.9209
1.250 0.1956 0.01044 0.00444 0.0333 0.3742 0.9231
1.500 0.2238 0.01049 0.00442 0.0332 0.3705 0.9238
1.750 0.2524 0.01051 0.00440 0.0330 0.3670 0.9245
2.000 0.2808 0.01055 0.00440 0.0328 0.3634 0.9252
2.250 0.3086 0.01059 0.00440 0.0328 0.3600 0.9259
2.500 0.3360 0.01065 0.00441 0.0327 0.3566 0.9268
2.750 0.3635 0.01070 0.00443 0.0327 0.3534 0.9277
3.000 0.3906 0.01073 0.00444 0.0327 0.3504 0.9286
3.250 0.4173 0.01076 0.00446 0.0329 0.3474 0.9296
3.500 0.4436 0.01082 0.00449 0.0330 0.3444 0.9308
3.750 0.4691 0.01088 0.00451 0.0333 0.3413 0.9320
4.000 0.4944 0.01092 0.00455 0.0337 0.3384 0.9331
4.250 0.5199 0.01094 0.00457 0.0340 0.3356 0.9342
4.500 0.5450 0.01097 0.00460 0.0343 0.3324 0.9355
4.750 0.5697 0.01102 0.00464 0.0348 0.3292 0.9369
5.000 0.5952 0.01109 0.00470 0.0350 0.3264 0.9378
5.250 0.6220 0.01118 0.00478 0.0350 0.3236 0.9383
5.500 0.6494 0.01123 0.00486 0.0349 0.3207 0.9388
5.750 0.6765 0.01130 0.00496 0.0348 0.3174 0.9393
6.000 0.7033 0.01139 0.00506 0.0347 0.3140 0.9399
6.250 0.7296 0.01151 0.00516 0.0348 0.3108 0.9406
6.500 0.7560 0.01162 0.00530 0.0348 0.3079 0.9412
6.750 0.7828 0.01170 0.00543 0.0347 0.3045 0.9421
7.000 0.8091 0.01180 0.00556 0.0348 0.3006 0.9430
7.250 0.8350 0.01193 0.00570 0.0348 0.2968 0.9438
7.500 0.8605 0.01207 0.00585 0.0349 0.2932 0.9446
7.750 0.8869 0.01216 0.00600 0.0349 0.2894 0.9455
8.000 0.9127 0.01228 0.00616 0.0349 0.2851 0.9463
8.250 0.9378 0.01244 0.00633 0.0350 0.2804 0.9473
8.500 0.9636 0.01258 0.00651 0.0350 0.2759 0.9482
8.750 0.9893 0.01272 0.00671 0.0350 0.2709 0.9491
9.000 1.0140 0.01293 0.00691 0.0351 0.2651 0.9500
9.250 1.0399 0.01309 0.00713 0.0350 0.2591 0.9508
9.500 1.0644 0.01331 0.00738 0.0351 0.2525 0.9516
9.750 1.0887 0.01352 0.00764 0.0353 0.2460 0.9524
10.000 1.1117 0.01380 0.00793 0.0356 0.2377 0.9534
10.250 1.1350 0.01407 0.00824 0.0359 0.2296 0.9545
10.500 1.1570 0.01442 0.00860 0.0363 0.2212 0.9557
10.750 1.1785 0.01480 0.00900 0.0367 0.2107 0.9569
11.000 1.1994 0.01521 0.00943 0.0372 0.2003 0.9582
11.250 1.2189 0.01570 0.00993 0.0378 0.1896 0.9597
11.500 1.2368 0.01630 0.01052 0.0385 0.1772 0.9613
11.750 1.2522 0.01697 0.01118 0.0395 0.1646 0.9630
12.000 1.2629 0.01768 0.01189 0.0412 0.1537 0.9650
12.250 1.2730 0.01856 0.01277 0.0428 0.1422 0.9676
12.500 1.2822 0.01958 0.01382 0.0440 0.1313 0.9705
12.750 1.2894 0.02083 0.01509 0.0450 0.1214 0.9735
13.000 1.2944 0.02241 0.01670 0.0456 0.1125 0.9774
13.250 1.2986 0.02443 0.01876 0.0454 0.1042 0.9816
13.500 1.3038 0.02663 0.02101 0.0446 0.0964 0.9869
13.750 1.3060 0.02939 0.02382 0.0433 0.0893 0.9922
14.000 1.3070 0.03254 0.02703 0.0415 0.0824 0.9962
14.250 1.3044 0.03572 0.03028 0.0401 0.0776 1.0000
14.500 1.2993 0.03919 0.03382 0.0390 0.0729 1.0000
14.750 1.2915 0.04303 0.03773 0.0377 0.0692 1.0000
15.000 1.2855 0.04666 0.04143 0.0365 0.0654 1.0000
15.250 1.2752 0.05080 0.04563 0.0352 0.0619 1.0000
15.500 1.2648 0.05496 0.04986 0.0339 0.0589 1.0000
15.750 1.2569 0.05895 0.05392 0.0326 0.0556 1.0000
16.000 1.2467 0.06331 0.05833 0.0312 0.0524 1.0000
16.250 1.2377 0.06761 0.06269 0.0297 0.0496 1.0000
16.500 1.2303 0.07173 0.06688 0.0282 0.0468 1.0000
16.750 1.2213 0.07613 0.07133 0.0267 0.0442 1.0000
17.000 1.2135 0.08044 0.07569 0.0252 0.0418 1.0000
17.250 1.2085 0.08437 0.07969 0.0237 0.0395 1.0000
17.500 1.2008 0.08872 0.08410 0.0221 0.0373 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 341 AIRFOIL (e341-il)