EPPLER 604 AIRFOIL (e604-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 604 AIRFOIL (e604-il) Reynolds number: 200,000 Max Cl/Cd: 67.12 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e604-il-200000-n5.txt Download as CSV file: xf-e604-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 604 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
------ -------- --------- --------- -------- -------- --------
-14.250 -0.4772 0.08413 0.08017 -0.0888 0.9932 0.0103
-14.000 -0.5009 0.07350 0.06929 -0.0973 0.9874 0.0101
-13.750 -0.5220 0.06485 0.06035 -0.1043 0.9797 0.0101
-13.500 -0.5364 0.05855 0.05378 -0.1087 0.9690 0.0100
-13.250 -0.5441 0.05351 0.04847 -0.1123 0.9583 0.0101
-13.000 -0.5452 0.04925 0.04395 -0.1154 0.9481 0.0101
-12.750 -0.5418 0.04505 0.03940 -0.1189 0.9399 0.0104
-12.500 -0.5310 0.04162 0.03563 -0.1220 0.9315 0.0106
-12.250 -0.5151 0.03870 0.03238 -0.1249 0.9231 0.0109
-12.000 -0.4944 0.03647 0.02999 -0.1277 0.9144 0.0111
-11.750 -0.4730 0.03459 0.02794 -0.1300 0.9047 0.0113
-11.250 -0.4314 0.03156 0.02459 -0.1334 0.8838 0.0122
-11.000 -0.4120 0.03038 0.02319 -0.1341 0.8735 0.0133
-10.750 -0.3949 0.02923 0.02192 -0.1343 0.8637 0.0142
-10.500 -0.3801 0.02827 0.02085 -0.1338 0.8542 0.0146
-10.250 -0.3637 0.02731 0.01976 -0.1334 0.8463 0.0152
-10.000 -0.3503 0.02638 0.01871 -0.1323 0.8379 0.0162
-9.750 -0.3362 0.02546 0.01766 -0.1313 0.8307 0.0170
-9.500 -0.3237 0.02467 0.01682 -0.1301 0.8231 0.0177
-9.250 -0.3107 0.02388 0.01595 -0.1288 0.8161 0.0188
-9.000 -0.2967 0.02317 0.01512 -0.1276 0.8099 0.0207
-8.750 -0.2844 0.02248 0.01440 -0.1261 0.8033 0.0223
-8.500 -0.2702 0.02183 0.01366 -0.1249 0.7980 0.0250
-8.250 -0.2578 0.02117 0.01298 -0.1233 0.7925 0.0274
-8.000 -0.2465 0.02053 0.01229 -0.1215 0.7867 0.0304
-7.750 -0.2330 0.01999 0.01167 -0.1199 0.7817 0.0346
-7.500 -0.2231 0.01947 0.01114 -0.1177 0.7766 0.0385
-7.250 -0.2140 0.01901 0.01067 -0.1154 0.7711 0.0437
-7.000 -0.2002 0.01852 0.01016 -0.1138 0.7668 0.0508
-6.500 -0.1701 0.01759 0.00924 -0.1109 0.7585 0.0724
-6.250 -0.1542 0.01714 0.00883 -0.1096 0.7540 0.0879
-6.000 -0.1365 0.01666 0.00839 -0.1086 0.7501 0.1088
-5.500 -0.1018 0.01559 0.00757 -0.1066 0.7428 0.1794
-5.250 -0.0859 0.01486 0.00714 -0.1056 0.7388 0.2452
-5.000 -0.0699 0.01390 0.00661 -0.1049 0.7352 0.3475
-4.750 -0.0515 0.01318 0.00659 -0.1040 0.7321 0.4976
-4.500 -0.0253 0.01333 0.00683 -0.1036 0.7289 0.5584
-4.250 0.0009 0.01350 0.00698 -0.1032 0.7253 0.5834
-4.000 0.0281 0.01365 0.00704 -0.1031 0.7220 0.6002
-3.750 0.0560 0.01379 0.00709 -0.1031 0.7188 0.6133
-3.500 0.0849 0.01392 0.00707 -0.1034 0.7160 0.6257
-3.250 0.1127 0.01423 0.00733 -0.1031 0.7133 0.6368
-3.000 0.1378 0.01460 0.00770 -0.1023 0.7099 0.6477
-2.750 0.1647 0.01479 0.00781 -0.1022 0.7065 0.6584
-2.500 0.1909 0.01505 0.00805 -0.1015 0.7033 0.6630
-2.250 0.2190 0.01515 0.00807 -0.1016 0.7006 0.6675
-2.000 0.2490 0.01510 0.00788 -0.1025 0.6983 0.6719
-1.750 0.2755 0.01506 0.00778 -0.1027 0.6949 0.6751
-1.500 0.3013 0.01509 0.00780 -0.1025 0.6913 0.6768
-1.250 0.3285 0.01511 0.00777 -0.1025 0.6882 0.6785
-1.000 0.3566 0.01512 0.00772 -0.1028 0.6854 0.6806
-0.750 0.3859 0.01511 0.00764 -0.1034 0.6829 0.6828
-0.500 0.4135 0.01511 0.00759 -0.1037 0.6799 0.6851
-0.250 0.4397 0.01511 0.00758 -0.1038 0.6764 0.6878
0.000 0.4680 0.01509 0.00749 -0.1044 0.6730 0.6909
0.250 0.4956 0.01510 0.00749 -0.1046 0.6700 0.6924
0.500 0.5242 0.01512 0.00747 -0.1050 0.6674 0.6939
0.750 0.5525 0.01516 0.00749 -0.1053 0.6647 0.6954
1.000 0.5765 0.01524 0.00760 -0.1049 0.6607 0.6972
1.250 0.6027 0.01529 0.00766 -0.1049 0.6572 0.6994
1.500 0.6307 0.01532 0.00768 -0.1053 0.6539 0.7018
1.750 0.6604 0.01533 0.00764 -0.1060 0.6511 0.7043
2.000 0.6889 0.01537 0.00766 -0.1065 0.6480 0.7067
2.250 0.7126 0.01546 0.00780 -0.1061 0.6437 0.7087
2.500 0.7381 0.01552 0.00790 -0.1059 0.6399 0.7102
2.750 0.7656 0.01556 0.00795 -0.1061 0.6365 0.7118
3.000 0.7953 0.01560 0.00798 -0.1067 0.6337 0.7139
3.250 0.8169 0.01574 0.00820 -0.1059 0.6290 0.7162
3.500 0.8414 0.01583 0.00832 -0.1056 0.6245 0.7187
3.750 0.8690 0.01586 0.00837 -0.1059 0.6206 0.7211
4.000 0.8996 0.01588 0.00835 -0.1067 0.6171 0.7235
4.250 0.9188 0.01602 0.00860 -0.1055 0.6112 0.7256
4.500 0.9426 0.01608 0.00871 -0.1049 0.6060 0.7273
4.750 0.9711 0.01608 0.00872 -0.1052 0.6018 0.7292
5.000 0.9889 0.01625 0.00900 -0.1037 0.5954 0.7317
5.250 1.0118 0.01632 0.00912 -0.1030 0.5895 0.7344
5.500 1.0377 0.01636 0.00918 -0.1029 0.5841 0.7372
5.750 1.0553 0.01652 0.00942 -0.1014 0.5766 0.7401
6.000 1.0805 0.01654 0.00946 -0.1012 0.5705 0.7423
6.250 1.0948 0.01672 0.00975 -0.0989 0.5628 0.7444
6.500 1.1146 0.01679 0.00985 -0.0976 0.5557 0.7468
6.750 1.1278 0.01697 0.01013 -0.0953 0.5473 0.7495
7.000 1.1471 0.01709 0.01027 -0.0939 0.5391 0.7525
7.250 1.1599 0.01739 0.01064 -0.0916 0.5292 0.7560
7.500 1.1762 0.01763 0.01090 -0.0899 0.5196 0.7590
7.750 1.1900 0.01792 0.01125 -0.0878 0.5089 0.7617
8.000 1.2017 0.01830 0.01169 -0.0854 0.4969 0.7647
8.250 1.2140 0.01872 0.01213 -0.0833 0.4843 0.7679
8.500 1.2257 0.01920 0.01262 -0.0811 0.4706 0.7714
8.750 1.2359 0.01978 0.01320 -0.0788 0.4558 0.7748
9.000 1.2436 0.02045 0.01388 -0.0762 0.4401 0.7777
9.250 1.2505 0.02123 0.01466 -0.0736 0.4238 0.7813
9.500 1.2562 0.02214 0.01556 -0.0711 0.4064 0.7854
9.750 1.2614 0.02318 0.01656 -0.0687 0.3888 0.7897
10.000 1.2639 0.02435 0.01772 -0.0660 0.3703 0.7933
10.250 1.2658 0.02564 0.01899 -0.0634 0.3523 0.7971
10.500 1.2675 0.02706 0.02038 -0.0610 0.3346 0.8012
10.750 1.2693 0.02859 0.02187 -0.0589 0.3166 0.8057
11.000 1.2706 0.03019 0.02346 -0.0567 0.2992 0.8096
11.250 1.2716 0.03188 0.02513 -0.0547 0.2822 0.8140
11.500 1.2730 0.03366 0.02691 -0.0529 0.2656 0.8191
11.750 1.2746 0.03551 0.02874 -0.0513 0.2497 0.8242
12.000 1.2762 0.03738 0.03061 -0.0498 0.2349 0.8292
12.250 1.2781 0.03934 0.03256 -0.0484 0.2204 0.8350
12.500 1.2802 0.04135 0.03457 -0.0472 0.2067 0.8408
12.750 1.2817 0.04344 0.03666 -0.0460 0.1934 0.8469
13.000 1.2841 0.04557 0.03880 -0.0450 0.1805 0.8540
13.250 1.2872 0.04759 0.04087 -0.0440 0.1683 0.8618
13.750 1.2917 0.05193 0.04529 -0.0423 0.1465 0.8837
14.500 1.2959 0.05852 0.05204 -0.0399 0.1192 1.0000
14.750 1.2985 0.06121 0.05473 -0.0399 0.1106 1.0000
15.000 1.3031 0.06372 0.05728 -0.0399 0.1024 1.0000
15.250 1.3051 0.06656 0.06013 -0.0400 0.0951 1.0000
15.500 1.3085 0.06927 0.06289 -0.0402 0.0881 1.0000
15.750 1.3109 0.07215 0.06580 -0.0405 0.0819 1.0000
16.000 1.3129 0.07513 0.06882 -0.0408 0.0759 1.0000
16.250 1.3149 0.07817 0.07190 -0.0412 0.0708 1.0000
16.500 1.3167 0.08126 0.07505 -0.0417 0.0657 1.0000
16.750 1.3169 0.08460 0.07845 -0.0423 0.0615 1.0000
17.000 1.3189 0.08774 0.08167 -0.0429 0.0572 1.0000
17.250 1.3169 0.09150 0.08547 -0.0438 0.0535 1.0000
17.500 1.3189 0.09473 0.08880 -0.0446 0.0499 1.0000
17.750 1.3180 0.09842 0.09256 -0.0456 0.0467 1.0000
18.000 1.3164 0.10225 0.09645 -0.0468 0.0440 1.0000
18.250 1.3165 0.10585 0.10017 -0.0479 0.0411 1.0000
18.500 1.3140 0.10990 0.10427 -0.0493 0.0388 1.0000
18.750 1.3121 0.11389 0.10835 -0.0508 0.0366 1.0000
19.000 1.3112 0.11775 0.11234 -0.0523 0.0344 1.0000
19.250 1.3081 0.12200 0.11666 -0.0541 0.0326 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 604 AIRFOIL (e604-il)