EPPLER 583 AIRFOIL (e583-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 583 AIRFOIL (e583-il) Reynolds number: 100,000 Max Cl/Cd: 32.62 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e583-il-100000-n5.txt Download as CSV file: xf-e583-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 583 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.1571 0.11934 0.11442 -0.0834 0.9264 0.0489
-11.500 -0.1465 0.11441 0.10946 -0.0893 0.9170 0.0496
-11.000 -0.1330 0.10345 0.09843 -0.1027 0.8976 0.0503
-10.500 -0.1080 0.08996 0.08476 -0.1080 0.8803 0.0253
-10.250 -0.1000 0.08596 0.08072 -0.1109 0.8696 0.0248
-10.000 -0.0967 0.08130 0.07601 -0.1143 0.8597 0.0242
-9.750 -0.0974 0.07616 0.07084 -0.1180 0.8503 0.0236
-9.500 -0.1365 0.06276 0.05731 -0.1279 0.8398 0.0215
-9.250 -0.1435 0.05920 0.05362 -0.1295 0.8315 0.0213
-9.000 -0.1574 0.05587 0.05017 -0.1294 0.8222 0.0212
-8.750 -0.1682 0.05312 0.04725 -0.1284 0.8153 0.0210
-8.500 -0.1821 0.05070 0.04468 -0.1259 0.8077 0.0209
-8.250 -0.1920 0.04771 0.04142 -0.1236 0.8014 0.0207
-8.000 -0.1954 0.04497 0.03838 -0.1215 0.7960 0.0206
-7.750 -0.1976 0.04230 0.03540 -0.1189 0.7897 0.0205
-7.500 -0.1940 0.03963 0.03233 -0.1167 0.7847 0.0204
-7.250 -0.1860 0.03695 0.02919 -0.1147 0.7808 0.0204
-7.000 -0.1771 0.03473 0.02659 -0.1124 0.7757 0.0206
-6.750 -0.1627 0.03264 0.02409 -0.1107 0.7713 0.0208
-6.500 -0.1437 0.03063 0.02162 -0.1095 0.7677 0.0212
-6.250 -0.1206 0.02893 0.01959 -0.1088 0.7647 0.0217
-6.000 -0.1019 0.02793 0.01850 -0.1076 0.7600 0.0225
-5.750 -0.0806 0.02702 0.01745 -0.1067 0.7558 0.0240
-5.500 -0.0573 0.02602 0.01625 -0.1060 0.7525 0.0263
-5.250 -0.0332 0.02521 0.01535 -0.1056 0.7497 0.0288
-5.000 -0.0108 0.02432 0.01434 -0.1046 0.7466 0.0310
-4.750 0.0068 0.02377 0.01380 -0.1031 0.7422 0.0340
-4.500 0.0268 0.02322 0.01322 -0.1020 0.7383 0.0395
-4.250 0.0479 0.02259 0.01255 -0.1009 0.7352 0.0468
-4.000 0.0704 0.02197 0.01192 -0.1001 0.7327 0.0610
-3.750 0.0893 0.02145 0.01153 -0.0988 0.7294 0.0861
-3.500 0.1050 0.02101 0.01135 -0.0970 0.7251 0.1313
-3.250 0.1214 0.02032 0.01114 -0.0955 0.7215 0.2288
-3.000 0.1325 0.01921 0.01092 -0.0931 0.7185 0.4279
-2.750 0.1463 0.01938 0.01215 -0.0876 0.7163 0.6929
-2.500 0.1617 0.01987 0.01252 -0.0847 0.7127 0.7517
-2.250 0.1749 0.02047 0.01303 -0.0814 0.7083 0.7826
-2.000 0.1920 0.02092 0.01334 -0.0786 0.7049 0.8075
-1.750 0.2145 0.02132 0.01359 -0.0765 0.7021 0.8264
-1.500 0.2407 0.02174 0.01387 -0.0746 0.6999 0.8488
-1.250 0.2619 0.02223 0.01426 -0.0720 0.6968 0.8730
-1.000 0.2836 0.02265 0.01462 -0.0704 0.6924 0.8894
-0.750 0.3108 0.02275 0.01460 -0.0704 0.6889 0.8961
-0.500 0.3296 0.02272 0.01445 -0.0690 0.6859 0.9028
-0.250 0.3643 0.02265 0.01421 -0.0705 0.6837 0.9057
0.000 0.3868 0.02275 0.01423 -0.0699 0.6804 0.9100
0.250 0.3850 0.02307 0.01456 -0.0651 0.6747 0.9174
0.500 0.4128 0.02311 0.01451 -0.0656 0.6713 0.9201
0.750 0.4430 0.02306 0.01436 -0.0663 0.6687 0.9227
1.000 0.4746 0.02297 0.01415 -0.0673 0.6666 0.9254
1.250 0.4675 0.02349 0.01473 -0.0618 0.6600 0.9323
1.500 0.4912 0.02366 0.01486 -0.0617 0.6560 0.9352
1.750 0.5208 0.02364 0.01477 -0.0624 0.6531 0.9376
2.000 0.5534 0.02353 0.01457 -0.0635 0.6509 0.9395
2.250 0.5503 0.02413 0.01524 -0.0589 0.6444 0.9464
2.500 0.5722 0.02439 0.01549 -0.0586 0.6398 0.9495
2.750 0.6031 0.02435 0.01541 -0.0595 0.6369 0.9516
3.000 0.6378 0.02422 0.01522 -0.0610 0.6348 0.9534
3.250 0.6348 0.02510 0.01620 -0.0570 0.6269 0.9607
3.500 0.6620 0.02529 0.01639 -0.0575 0.6228 0.9638
3.750 0.6969 0.02517 0.01625 -0.0591 0.6202 0.9658
4.000 0.7353 0.02495 0.01600 -0.0611 0.6182 0.9673
4.250 0.7282 0.02627 0.01745 -0.0572 0.6084 0.9760
4.500 0.7634 0.02618 0.01736 -0.0588 0.6053 0.9786
4.750 0.8040 0.02592 0.01711 -0.0612 0.6031 0.9804
5.250 0.8344 0.02713 0.01846 -0.0596 0.5897 0.9978
5.500 0.8718 0.02673 0.01807 -0.0612 0.5873 1.0000
6.000 0.8531 0.02817 0.01959 -0.0510 0.5725 1.0000
6.250 0.8944 0.02752 0.01896 -0.0528 0.5706 1.0000
6.750 0.9098 0.02906 0.02060 -0.0490 0.5558 1.0000
7.250 0.9330 0.03064 0.02231 -0.0466 0.5409 1.0000
7.750 0.9607 0.03215 0.02394 -0.0449 0.5259 1.0000
8.250 0.9926 0.03343 0.02535 -0.0436 0.5108 1.0000
8.750 1.0272 0.03453 0.02661 -0.0425 0.4956 1.0000
9.250 1.0657 0.03532 0.02754 -0.0415 0.4801 1.0000
9.750 1.1080 0.03579 0.02817 -0.0408 0.4639 1.0000
10.000 1.1103 0.03756 0.03001 -0.0394 0.4519 1.0000
10.250 1.1295 0.03795 0.03047 -0.0389 0.4426 1.0000
10.500 1.1553 0.03775 0.03035 -0.0387 0.4337 1.0000
10.750 1.1623 0.03917 0.03185 -0.0376 0.4214 1.0000
11.000 1.1784 0.03983 0.03257 -0.0370 0.4103 1.0000
11.250 1.2034 0.03968 0.03247 -0.0367 0.3998 1.0000
11.500 1.2145 0.04075 0.03360 -0.0358 0.3871 1.0000
11.750 1.2222 0.04214 0.03504 -0.0348 0.3736 1.0000
12.000 1.2314 0.04342 0.03638 -0.0339 0.3603 1.0000
12.250 1.2405 0.04472 0.03772 -0.0331 0.3465 1.0000
12.500 1.2486 0.04613 0.03916 -0.0322 0.3322 1.0000
12.750 1.2560 0.04765 0.04071 -0.0314 0.3181 1.0000
13.000 1.2623 0.04930 0.04237 -0.0306 0.3041 1.0000
13.250 1.2675 0.05110 0.04418 -0.0298 0.2899 1.0000
13.500 1.2721 0.05301 0.04612 -0.0291 0.2761 1.0000
13.750 1.2754 0.05511 0.04822 -0.0285 0.2622 1.0000
14.000 1.2766 0.05751 0.05066 -0.0279 0.2485 1.0000
14.250 1.2772 0.06006 0.05324 -0.0275 0.2351 1.0000
14.500 1.2774 0.06274 0.05597 -0.0272 0.2219 1.0000
14.750 1.2771 0.06553 0.05880 -0.0269 0.2091 1.0000
15.000 1.2762 0.06845 0.06174 -0.0268 0.1965 1.0000
15.250 1.2747 0.07152 0.06485 -0.0268 0.1844 1.0000
15.500 1.2725 0.07474 0.06808 -0.0269 0.1728 1.0000
15.750 1.2697 0.07816 0.07154 -0.0272 0.1614 1.0000
16.000 1.2668 0.08169 0.07514 -0.0276 0.1504 1.0000
16.250 1.2636 0.08531 0.07881 -0.0282 0.1402 1.0000
16.500 1.2593 0.08913 0.08265 -0.0288 0.1305 1.0000
16.750 1.2551 0.09307 0.08664 -0.0297 0.1211 1.0000
17.000 1.2514 0.09701 0.09067 -0.0306 0.1123 1.0000
17.250 1.2465 0.10114 0.09484 -0.0317 0.1044 1.0000
17.500 1.2419 0.10533 0.09910 -0.0330 0.0967 1.0000
17.750 1.2379 0.10951 0.10336 -0.0343 0.0896 1.0000
18.000 1.2323 0.11396 0.10783 -0.0359 0.0834 1.0000
18.250 1.2288 0.11816 0.11215 -0.0374 0.0770 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 583 AIRFOIL (e583-il)