EPPLER 642 AIRFOIL (e642-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 642 AIRFOIL (e642-il) Reynolds number: 200,000 Max Cl/Cd: 77.07 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e642-il-200000.txt Download as CSV file: xf-e642-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 642 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.2578 0.09550 0.09204 -0.0766 0.9030 0.0560
-10.500 -0.2732 0.08755 0.08402 -0.0881 0.8844 0.0587
-10.250 -0.2992 0.08037 0.07670 -0.0965 0.8696 0.0590
-10.000 -0.3200 0.07669 0.07287 -0.0982 0.8586 0.0591
-9.750 -0.3470 0.07413 0.07012 -0.0970 0.8494 0.0592
-9.500 -0.3113 0.06764 0.06381 -0.0986 0.8463 0.0612
-9.250 -0.3094 0.06514 0.06127 -0.0981 0.8399 0.0622
-9.000 -0.3142 0.06258 0.05864 -0.0972 0.8342 0.0635
-8.750 -0.3260 0.06008 0.05607 -0.0954 0.8278 0.0645
-8.500 -0.3363 0.05773 0.05359 -0.0933 0.8229 0.0660
-8.250 -0.3540 0.05560 0.05113 -0.0903 0.8184 0.0702
-8.000 -0.3825 0.05412 0.04908 -0.0852 0.8134 0.0723
-7.750 -0.3573 0.04934 0.04461 -0.0868 0.8105 0.0749
-7.500 -0.3444 0.04745 0.04266 -0.0859 0.8073 0.0780
-6.750 -0.3387 0.03372 0.02703 -0.0750 0.7971 0.0497
-6.500 -0.3173 0.03007 0.02271 -0.0726 0.7944 0.0384
-6.250 -0.2957 0.02724 0.01964 -0.0718 0.7920 0.0372
-6.000 -0.2716 0.02531 0.01742 -0.0711 0.7898 0.0365
-5.750 -0.2459 0.02379 0.01567 -0.0706 0.7872 0.0363
-5.500 -0.2193 0.02254 0.01425 -0.0704 0.7841 0.0366
-5.250 -0.1927 0.02167 0.01326 -0.0701 0.7814 0.0377
-5.000 -0.1661 0.02073 0.01224 -0.0699 0.7790 0.0394
-4.750 -0.1410 0.01970 0.01124 -0.0696 0.7766 0.0411
-4.500 -0.1175 0.01906 0.01058 -0.0688 0.7744 0.0429
-4.250 -0.0956 0.01857 0.01008 -0.0679 0.7719 0.0457
-4.000 -0.0755 0.01811 0.00961 -0.0667 0.7690 0.0492
-3.750 -0.0556 0.01766 0.00918 -0.0655 0.7661 0.0580
-3.500 -0.0427 0.01652 0.00859 -0.0631 0.7632 0.1471
-3.250 -0.0387 0.01511 0.00831 -0.0597 0.7607 0.3930
-3.000 -0.0285 0.01440 0.00910 -0.0549 0.7586 0.7041
-2.750 -0.0090 0.01562 0.01038 -0.0515 0.7559 0.8091
-2.500 0.0088 0.01669 0.01139 -0.0482 0.7527 0.8431
-2.250 0.0608 0.01811 0.01270 -0.0499 0.7504 0.8678
-2.000 0.1200 0.01876 0.01316 -0.0545 0.7482 0.8823
-1.750 0.1886 0.01910 0.01333 -0.0616 0.7463 0.8946
-1.500 0.2530 0.01935 0.01340 -0.0680 0.7445 0.9095
-1.250 0.2962 0.01944 0.01336 -0.0708 0.7425 0.9211
-1.000 0.3248 0.01965 0.01355 -0.0714 0.7391 0.9323
-0.750 0.3749 0.01948 0.01332 -0.0762 0.7359 0.9383
-0.500 0.4034 0.01951 0.01330 -0.0767 0.7327 0.9471
-0.250 0.4466 0.01927 0.01298 -0.0802 0.7301 0.9524
0.000 0.4733 0.01931 0.01295 -0.0802 0.7275 0.9599
0.250 0.5139 0.01913 0.01275 -0.0834 0.7242 0.9645
0.500 0.5423 0.01918 0.01282 -0.0843 0.7198 0.9704
0.750 0.5761 0.01902 0.01263 -0.0860 0.7163 0.9749
1.000 0.6109 0.01883 0.01239 -0.0878 0.7133 0.9790
1.250 0.6405 0.01880 0.01230 -0.0885 0.7107 0.9832
1.500 0.6712 0.01881 0.01240 -0.0901 0.7050 0.9869
1.750 0.7035 0.01867 0.01227 -0.0915 0.7008 0.9904
2.000 0.7345 0.01851 0.01207 -0.0925 0.6974 0.9935
2.250 0.7672 0.01837 0.01193 -0.0939 0.6936 0.9963
2.500 0.7956 0.01839 0.01203 -0.0948 0.6874 0.9990
2.750 0.8226 0.01825 0.01189 -0.0949 0.6832 1.0000
3.000 0.8472 0.01812 0.01172 -0.0944 0.6799 1.0000
3.250 0.8637 0.01838 0.01208 -0.0929 0.6729 1.0000
3.500 0.8865 0.01827 0.01200 -0.0920 0.6680 1.0000
3.750 0.9122 0.01804 0.01173 -0.0916 0.6644 1.0000
4.000 0.9285 0.01825 0.01205 -0.0899 0.6568 1.0000
4.250 0.9522 0.01803 0.01184 -0.0891 0.6517 1.0000
4.500 0.9767 0.01783 0.01164 -0.0885 0.6470 1.0000
4.750 0.9941 0.01787 0.01179 -0.0868 0.6390 1.0000
5.000 1.0206 0.01750 0.01140 -0.0864 0.6343 1.0000
5.250 1.0374 0.01756 0.01159 -0.0847 0.6262 1.0000
5.500 1.0620 0.01722 0.01126 -0.0840 0.6202 1.0000
5.750 1.0807 0.01715 0.01129 -0.0824 0.6124 1.0000
6.000 1.1033 0.01688 0.01106 -0.0814 0.6054 1.0000
6.250 1.1215 0.01677 0.01106 -0.0797 0.5967 1.0000
6.500 1.1449 0.01641 0.01071 -0.0787 0.5889 1.0000
6.750 1.1597 0.01636 0.01079 -0.0764 0.5779 1.0000
7.000 1.1768 0.01621 0.01074 -0.0744 0.5669 1.0000
7.250 1.1939 0.01605 0.01065 -0.0724 0.5549 1.0000
7.500 1.2086 0.01597 0.01062 -0.0700 0.5403 1.0000
7.750 1.2215 0.01595 0.01065 -0.0672 0.5237 1.0000
8.000 1.2323 0.01599 0.01068 -0.0641 0.5043 1.0000
8.250 1.2392 0.01618 0.01080 -0.0603 0.4824 1.0000
8.500 1.2413 0.01650 0.01105 -0.0557 0.4597 1.0000
8.750 1.2384 0.01690 0.01138 -0.0503 0.4376 1.0000
9.000 1.2283 0.01735 0.01176 -0.0436 0.4167 1.0000
9.250 1.2076 0.01768 0.01205 -0.0350 0.4015 1.0000
9.500 1.1854 0.01803 0.01236 -0.0264 0.3879 1.0000
9.750 1.1644 0.01845 0.01274 -0.0185 0.3737 1.0000
10.000 1.1523 0.01919 0.01340 -0.0128 0.3546 1.0000
10.250 1.1486 0.02030 0.01441 -0.0094 0.3312 1.0000
10.500 1.1466 0.02169 0.01569 -0.0068 0.3070 1.0000
10.750 1.1468 0.02321 0.01715 -0.0048 0.2826 1.0000
11.000 1.1471 0.02487 0.01870 -0.0030 0.2610 1.0000
11.250 1.1488 0.02655 0.02032 -0.0016 0.2382 1.0000
11.500 1.1488 0.02843 0.02209 -0.0002 0.2175 1.0000
11.750 1.1518 0.03018 0.02380 0.0009 0.1974 1.0000
12.000 1.1543 0.03201 0.02558 0.0020 0.1793 1.0000
12.250 1.1565 0.03392 0.02744 0.0030 0.1624 1.0000
12.500 1.1580 0.03592 0.02939 0.0039 0.1459 1.0000
12.750 1.1611 0.03789 0.03135 0.0047 0.1287 1.0000
13.000 1.1620 0.04010 0.03350 0.0054 0.1103 1.0000
13.250 1.1595 0.04269 0.03598 0.0062 0.0887 1.0000
13.500 1.1545 0.04563 0.03877 0.0069 0.0697 1.0000
13.750 1.1457 0.04905 0.04206 0.0077 0.0511 1.0000
14.000 1.1371 0.05263 0.04559 0.0083 0.0413 1.0000
14.250 1.1313 0.05606 0.04907 0.0086 0.0359 1.0000
14.500 1.1277 0.05936 0.05241 0.0086 0.0331 1.0000
14.750 1.1224 0.06292 0.05601 0.0087 0.0312 1.0000
15.000 1.1238 0.06583 0.05904 0.0086 0.0292 1.0000
15.250 1.1252 0.06873 0.06202 0.0084 0.0281 1.0000
15.500 1.1264 0.07171 0.06507 0.0081 0.0269 1.0000
15.750 1.1277 0.07462 0.06800 0.0079 0.0260 1.0000
16.000 1.1316 0.07715 0.07057 0.0081 0.0248 1.0000
16.250 1.1366 0.07982 0.07338 0.0077 0.0239 1.0000
16.500 1.1423 0.08235 0.07602 0.0074 0.0231 1.0000
16.750 1.1471 0.08506 0.07883 0.0070 0.0223 1.0000
17.000 1.1529 0.08761 0.08147 0.0066 0.0217 1.0000
17.250 1.1591 0.09014 0.08408 0.0063 0.0212 1.0000
17.500 1.1662 0.09251 0.08651 0.0060 0.0207 1.0000
17.750 1.1745 0.09474 0.08882 0.0061 0.0203 1.0000
18.000 1.1803 0.09752 0.09175 0.0058 0.0201 1.0000
18.250 1.1818 0.10111 0.09551 0.0053 0.0199 1.0000
18.500 1.1768 0.10561 0.10022 0.0036 0.0198 1.0000
18.750 1.1698 0.11054 0.10537 0.0016 0.0198 1.0000
19.000 1.1596 0.11607 0.11112 -0.0009 0.0197 1.0000
19.250 1.1507 0.12154 0.11679 -0.0034 0.0198 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 642 AIRFOIL (e642-il)