EPPLER 420 AIRFOIL (e420-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 420 AIRFOIL (e420-il) Reynolds number: 1,000,000 Max Cl/Cd: 134.13 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e420-il-1000000-n5.txt Download as CSV file: xf-e420-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 420 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 0.4471 0.07296 0.06865 -0.1687 0.5277 0.0106
-7.500 0.4649 0.07023 0.06595 -0.1686 0.5221 0.0117
-7.000 0.2965 0.03254 0.02803 -0.1940 0.5253 0.0082
-6.750 0.3650 0.01917 0.01396 -0.2226 0.5223 0.0082
-6.500 0.4007 0.01615 0.01064 -0.2274 0.5186 0.0083
-6.250 0.4306 0.01513 0.00945 -0.2289 0.5147 0.0083
-5.750 0.4894 0.01363 0.00765 -0.2307 0.5085 0.0085
-5.500 0.5190 0.01297 0.00685 -0.2315 0.5061 0.0087
-5.250 0.5475 0.01256 0.00634 -0.2318 0.5036 0.0089
-5.000 0.5755 0.01224 0.00595 -0.2319 0.5008 0.0092
-4.750 0.6031 0.01200 0.00563 -0.2319 0.4981 0.0094
-4.500 0.6304 0.01180 0.00535 -0.2318 0.4953 0.0097
-4.000 0.6852 0.01145 0.00488 -0.2316 0.4905 0.0103
-3.750 0.7129 0.01129 0.00467 -0.2315 0.4885 0.0106
-3.500 0.7404 0.01117 0.00449 -0.2314 0.4862 0.0108
-3.250 0.7679 0.01102 0.00430 -0.2313 0.4839 0.0113
-3.000 0.7950 0.01092 0.00415 -0.2311 0.4817 0.0118
-2.750 0.8218 0.01084 0.00404 -0.2308 0.4795 0.0125
-2.500 0.8483 0.01080 0.00395 -0.2305 0.4772 0.0133
-2.250 0.8748 0.01075 0.00386 -0.2301 0.4749 0.0146
-2.000 0.9016 0.01071 0.00380 -0.2299 0.4731 0.0164
-1.750 0.9288 0.01064 0.00373 -0.2297 0.4718 0.0192
-1.500 0.9560 0.01059 0.00368 -0.2295 0.4701 0.0234
-1.250 0.9830 0.01054 0.00365 -0.2293 0.4683 0.0313
-1.000 1.0098 0.01052 0.00364 -0.2291 0.4663 0.0404
-0.750 1.0363 0.01052 0.00365 -0.2288 0.4645 0.0499
-0.500 1.0627 0.01052 0.00368 -0.2285 0.4627 0.0614
-0.250 1.0888 0.01055 0.00373 -0.2282 0.4609 0.0730
0.250 1.1399 0.01067 0.00386 -0.2273 0.4567 0.0959
0.500 1.1659 0.01072 0.00392 -0.2269 0.4552 0.1053
0.750 1.1922 0.01075 0.00399 -0.2266 0.4540 0.1155
1.000 1.2182 0.01080 0.00406 -0.2263 0.4526 0.1243
1.250 1.2440 0.01087 0.00414 -0.2259 0.4512 0.1321
1.500 1.2696 0.01093 0.00423 -0.2255 0.4497 0.1419
1.750 1.2949 0.01102 0.00432 -0.2250 0.4482 0.1503
2.000 1.3200 0.01111 0.00443 -0.2245 0.4464 0.1602
2.250 1.3447 0.01121 0.00454 -0.2240 0.4445 0.1679
2.500 1.3690 0.01133 0.00466 -0.2234 0.4426 0.1769
2.750 1.3930 0.01147 0.00479 -0.2227 0.4407 0.1835
3.000 1.4166 0.01161 0.00495 -0.2220 0.4388 0.1922
3.250 1.4414 0.01171 0.00507 -0.2215 0.4379 0.2000
3.500 1.4662 0.01181 0.00520 -0.2210 0.4367 0.2079
3.750 1.4908 0.01191 0.00534 -0.2206 0.4353 0.2164
4.000 1.5151 0.01203 0.00547 -0.2200 0.4338 0.2229
4.250 1.5391 0.01215 0.00563 -0.2194 0.4321 0.2332
4.500 1.5629 0.01229 0.00579 -0.2188 0.4306 0.2409
4.750 1.5864 0.01243 0.00596 -0.2182 0.4290 0.2526
5.000 1.6092 0.01260 0.00615 -0.2174 0.4272 0.2621
5.250 1.6312 0.01278 0.00636 -0.2166 0.4252 0.2754
5.500 1.6524 0.01300 0.00660 -0.2156 0.4231 0.2889
5.750 1.6745 0.01319 0.00683 -0.2147 0.4215 0.3042
6.000 1.6982 0.01332 0.00703 -0.2142 0.4203 0.3273
6.250 1.7215 0.01346 0.00724 -0.2137 0.4188 0.3545
6.500 1.7446 0.01360 0.00748 -0.2131 0.4169 0.4010
6.750 1.7665 0.01317 0.00789 -0.2125 0.4148 1.0000
7.000 1.7874 0.01341 0.00814 -0.2115 0.4125 1.0000
7.250 1.8075 0.01369 0.00842 -0.2104 0.4101 1.0000
7.500 1.8266 0.01401 0.00873 -0.2092 0.4074 1.0000
7.750 1.8462 0.01432 0.00904 -0.2080 0.4049 1.0000
8.000 1.8676 0.01456 0.00931 -0.2073 0.4028 1.0000
8.250 1.8880 0.01483 0.00961 -0.2063 0.4001 1.0000
8.500 1.9074 0.01515 0.00994 -0.2052 0.3969 1.0000
8.750 1.9252 0.01554 0.01033 -0.2039 0.3936 1.0000
9.000 1.9408 0.01601 0.01081 -0.2022 0.3903 1.0000
9.250 1.9598 0.01637 0.01120 -0.2012 0.3877 1.0000
9.500 1.9785 0.01674 0.01160 -0.2001 0.3843 1.0000
9.750 1.9953 0.01720 0.01208 -0.1987 0.3807 1.0000
10.000 2.0102 0.01776 0.01264 -0.1972 0.3769 1.0000
10.250 2.0235 0.01840 0.01329 -0.1954 0.3731 1.0000
10.500 2.0406 0.01889 0.01382 -0.1942 0.3695 1.0000
10.750 2.0549 0.01952 0.01448 -0.1927 0.3649 1.0000
11.000 2.0652 0.02039 0.01535 -0.1907 0.3597 1.0000
11.250 2.0782 0.02113 0.01613 -0.1891 0.3555 1.0000
11.500 2.0897 0.02200 0.01701 -0.1874 0.3493 1.0000
11.750 2.0951 0.02326 0.01826 -0.1850 0.3431 1.0000
12.000 2.1064 0.02421 0.01925 -0.1834 0.3371 1.0000
12.250 2.1115 0.02560 0.02065 -0.1812 0.3302 1.0000
12.500 2.1161 0.02708 0.02215 -0.1791 0.3231 1.0000
12.750 2.1177 0.02886 0.02394 -0.1768 0.3150 1.0000
13.000 2.1189 0.03075 0.02584 -0.1746 0.3068 1.0000
13.250 2.1139 0.03324 0.02833 -0.1721 0.2979 1.0000
13.500 2.1127 0.03552 0.03063 -0.1701 0.2888 1.0000
13.750 2.1050 0.03848 0.03359 -0.1678 0.2802 1.0000
14.000 2.0954 0.04176 0.03690 -0.1656 0.2703 1.0000
14.250 2.0868 0.04508 0.04024 -0.1636 0.2619 1.0000
14.500 2.0706 0.04933 0.04450 -0.1616 0.2525 1.0000
14.750 2.0595 0.05319 0.04840 -0.1600 0.2435 1.0000
15.000 2.0436 0.05773 0.05296 -0.1584 0.2351 1.0000
15.250 2.0299 0.06215 0.05741 -0.1572 0.2270 1.0000
15.500 2.0151 0.06683 0.06213 -0.1561 0.2190 1.0000
15.750 1.9989 0.07182 0.06714 -0.1552 0.2109 1.0000
16.000 1.9889 0.07610 0.07146 -0.1545 0.2038 1.0000
16.250 1.9728 0.08128 0.07667 -0.1540 0.1962 1.0000
16.500 1.9629 0.08575 0.08117 -0.1536 0.1897 1.0000
16.750 1.9495 0.09072 0.08617 -0.1534 0.1823 1.0000
17.000 1.9424 0.09489 0.09037 -0.1533 0.1759 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 420 AIRFOIL (e420-il)