EPPLER 562 AIRFOIL (e562-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 562 AIRFOIL (e562-il) Reynolds number: 1,000,000 Max Cl/Cd: 115.05 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e562-il-1000000-n5.txt Download as CSV file: xf-e562-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 562 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
------ -------- --------- --------- -------- -------- --------
-17.000 -0.8138 0.10122 0.09827 -0.0502 1.0000 0.0059
-16.750 -0.8337 0.09440 0.09137 -0.0535 1.0000 0.0060
-16.500 -0.8457 0.08909 0.08597 -0.0559 1.0000 0.0059
-16.250 -0.8634 0.08298 0.07977 -0.0589 1.0000 0.0059
-16.000 -0.8801 0.07723 0.07394 -0.0616 1.0000 0.0060
-15.750 -0.8949 0.07191 0.06854 -0.0640 1.0000 0.0060
-15.500 -0.9124 0.06643 0.06297 -0.0665 1.0000 0.0061
-15.250 -0.9279 0.06150 0.05796 -0.0686 1.0000 0.0060
-15.000 -0.9457 0.05663 0.05301 -0.0704 1.0000 0.0060
-14.750 -0.9663 0.05164 0.04794 -0.0722 1.0000 0.0061
-14.500 -0.9832 0.04737 0.04357 -0.0734 1.0000 0.0061
-14.250 -0.9978 0.04283 0.03894 -0.0757 0.9996 0.0062
-14.000 -0.9922 0.03841 0.03442 -0.0814 0.9975 0.0063
-13.750 -0.9824 0.03437 0.03028 -0.0872 0.9950 0.0064
-13.500 -0.9740 0.03064 0.02644 -0.0923 0.9923 0.0065
-13.250 -0.9645 0.02728 0.02299 -0.0969 0.9880 0.0067
-13.000 -0.9371 0.02409 0.01966 -0.1046 0.9855 0.0068
-12.750 -0.9074 0.02240 0.01787 -0.1087 0.9841 0.0070
-12.500 -0.8875 0.02123 0.01663 -0.1096 0.9810 0.0072
-12.250 -0.8650 0.02020 0.01553 -0.1104 0.9777 0.0074
-12.000 -0.8370 0.01932 0.01458 -0.1120 0.9756 0.0077
-11.750 -0.8074 0.01842 0.01361 -0.1138 0.9740 0.0079
-11.500 -0.7763 0.01757 0.01272 -0.1158 0.9728 0.0084
-11.250 -0.7586 0.01691 0.01203 -0.1147 0.9677 0.0087
-11.000 -0.7316 0.01628 0.01136 -0.1153 0.9645 0.0092
-10.750 -0.7014 0.01567 0.01071 -0.1165 0.9622 0.0096
-10.500 -0.6696 0.01511 0.01009 -0.1180 0.9605 0.0101
-10.250 -0.6521 0.01459 0.00954 -0.1164 0.9526 0.0105
-10.000 -0.6227 0.01402 0.00895 -0.1172 0.9491 0.0111
-9.750 -0.5981 0.01357 0.00847 -0.1170 0.9427 0.0118
-9.500 -0.5685 0.01311 0.00796 -0.1176 0.9377 0.0125
-9.250 -0.5341 0.01266 0.00747 -0.1193 0.9342 0.0131
-9.000 -0.5019 0.01215 0.00694 -0.1205 0.9274 0.0141
-8.750 -0.4617 0.01169 0.00645 -0.1234 0.9223 0.0152
-8.500 -0.4176 0.01130 0.00600 -0.1270 0.9162 0.0163
-8.250 -0.3669 0.01083 0.00549 -0.1322 0.9080 0.0179
-8.000 -0.3170 0.01047 0.00506 -0.1372 0.8950 0.0195
-7.750 -0.2775 0.01021 0.00469 -0.1398 0.8748 0.0212
-7.500 -0.2472 0.01000 0.00438 -0.1405 0.8518 0.0236
-7.250 -0.2206 0.00981 0.00410 -0.1403 0.8296 0.0272
-6.750 -0.1702 0.00934 0.00354 -0.1396 0.7906 0.0467
-6.500 -0.1449 0.00911 0.00329 -0.1392 0.7724 0.0601
-6.250 -0.1194 0.00889 0.00305 -0.1388 0.7554 0.0749
-6.000 -0.0936 0.00869 0.00285 -0.1385 0.7392 0.0906
-5.750 -0.0673 0.00853 0.00266 -0.1383 0.7241 0.1055
-5.500 -0.0410 0.00836 0.00249 -0.1381 0.7094 0.1223
-5.250 -0.0145 0.00819 0.00233 -0.1379 0.6947 0.1427
-5.000 0.0124 0.00802 0.00218 -0.1378 0.6810 0.1672
-4.750 0.0396 0.00792 0.00207 -0.1376 0.6677 0.1832
-4.500 0.0669 0.00783 0.00198 -0.1375 0.6549 0.1998
-4.250 0.0942 0.00777 0.00190 -0.1374 0.6424 0.2145
-4.000 0.1215 0.00772 0.00183 -0.1373 0.6301 0.2289
-3.750 0.1492 0.00767 0.00177 -0.1372 0.6178 0.2448
-3.500 0.1769 0.00764 0.00172 -0.1371 0.6064 0.2545
-3.000 0.2321 0.00765 0.00165 -0.1369 0.5834 0.2705
-2.750 0.2600 0.00767 0.00162 -0.1368 0.5730 0.2775
-2.500 0.2877 0.00768 0.00160 -0.1367 0.5622 0.2858
-2.250 0.3155 0.00769 0.00158 -0.1366 0.5516 0.2949
-2.000 0.3434 0.00769 0.00158 -0.1366 0.5422 0.3070
-1.750 0.3710 0.00773 0.00158 -0.1364 0.5322 0.3147
-1.500 0.3990 0.00775 0.00158 -0.1364 0.5224 0.3215
-1.250 0.4267 0.00780 0.00158 -0.1363 0.5134 0.3278
-1.000 0.4545 0.00783 0.00160 -0.1362 0.5044 0.3360
-0.750 0.4823 0.00788 0.00162 -0.1361 0.4959 0.3428
-0.500 0.5099 0.00793 0.00165 -0.1360 0.4863 0.3500
-0.250 0.5377 0.00798 0.00167 -0.1359 0.4786 0.3565
0.000 0.5653 0.00803 0.00171 -0.1357 0.4701 0.3638
0.250 0.5930 0.00809 0.00175 -0.1356 0.4621 0.3718
0.500 0.6204 0.00815 0.00180 -0.1355 0.4539 0.3787
0.750 0.6481 0.00821 0.00185 -0.1354 0.4468 0.3864
1.000 0.6754 0.00829 0.00191 -0.1352 0.4384 0.3936
1.250 0.7029 0.00835 0.00198 -0.1351 0.4313 0.4012
1.500 0.7302 0.00844 0.00204 -0.1349 0.4231 0.4085
1.750 0.7575 0.00851 0.00212 -0.1347 0.4163 0.4160
2.000 0.7847 0.00860 0.00219 -0.1346 0.4090 0.4228
2.250 0.8118 0.00868 0.00228 -0.1344 0.4021 0.4305
2.500 0.8390 0.00876 0.00237 -0.1342 0.3954 0.4379
3.000 0.8929 0.00895 0.00256 -0.1337 0.3820 0.4528
3.250 0.9195 0.00906 0.00267 -0.1334 0.3752 0.4601
3.500 0.9463 0.00916 0.00278 -0.1332 0.3693 0.4682
3.750 0.9729 0.00927 0.00289 -0.1329 0.3622 0.4759
4.000 0.9991 0.00939 0.00302 -0.1326 0.3558 0.4849
4.250 1.0258 0.00949 0.00314 -0.1323 0.3495 0.4932
4.500 1.0516 0.00964 0.00329 -0.1319 0.3420 0.5031
4.750 1.0779 0.00975 0.00343 -0.1316 0.3354 0.5127
5.250 1.1292 0.01003 0.00374 -0.1307 0.3204 0.5354
5.500 1.1543 0.01020 0.00392 -0.1302 0.3116 0.5475
5.750 1.1797 0.01034 0.00409 -0.1298 0.3043 0.5615
6.000 1.2041 0.01054 0.00430 -0.1291 0.2951 0.5770
6.250 1.2291 0.01070 0.00449 -0.1286 0.2868 0.5937
6.500 1.2531 0.01091 0.00471 -0.1279 0.2776 0.6122
6.750 1.2771 0.01110 0.00494 -0.1272 0.2677 0.6350
7.000 1.3004 0.01132 0.00519 -0.1264 0.2566 0.6586
7.500 1.3447 0.01182 0.00576 -0.1244 0.2344 0.7213
7.750 1.3658 0.01204 0.00605 -0.1232 0.2231 0.7620
8.000 1.3854 0.01226 0.00637 -0.1217 0.2128 0.8175
8.250 1.3972 0.01232 0.00665 -0.1184 0.2022 0.9796
8.500 1.4153 0.01272 0.00700 -0.1167 0.1896 1.0000
8.750 1.4328 0.01319 0.00740 -0.1149 0.1748 1.0000
9.000 1.4485 0.01375 0.00786 -0.1129 0.1564 1.0000
9.250 1.4619 0.01440 0.00840 -0.1105 0.1371 1.0000
9.500 1.4764 0.01499 0.00892 -0.1083 0.1232 1.0000
9.750 1.4917 0.01554 0.00943 -0.1063 0.1122 1.0000
10.000 1.5058 0.01615 0.00998 -0.1042 0.1016 1.0000
10.250 1.5190 0.01679 0.01058 -0.1019 0.0917 1.0000
10.500 1.5307 0.01751 0.01125 -0.0995 0.0813 1.0000
10.750 1.5415 0.01828 0.01198 -0.0970 0.0715 1.0000
11.000 1.5521 0.01907 0.01274 -0.0946 0.0631 1.0000
11.250 1.5603 0.02001 0.01364 -0.0920 0.0539 1.0000
11.500 1.5697 0.02090 0.01452 -0.0896 0.0476 1.0000
11.750 1.5777 0.02191 0.01552 -0.0872 0.0420 1.0000
12.000 1.5873 0.02285 0.01648 -0.0851 0.0381 1.0000
12.250 1.5946 0.02398 0.01761 -0.0828 0.0342 1.0000
12.500 1.6023 0.02513 0.01879 -0.0808 0.0308 1.0000
12.750 1.6082 0.02645 0.02014 -0.0787 0.0275 1.0000
13.000 1.6139 0.02786 0.02157 -0.0768 0.0247 1.0000
13.250 1.6188 0.02941 0.02315 -0.0750 0.0221 1.0000
13.500 1.6224 0.03113 0.02490 -0.0733 0.0197 1.0000
13.750 1.6273 0.03284 0.02666 -0.0719 0.0183 1.0000
14.000 1.6311 0.03473 0.02859 -0.0706 0.0168 1.0000
14.250 1.6331 0.03687 0.03079 -0.0694 0.0154 1.0000
14.500 1.6367 0.03897 0.03295 -0.0685 0.0145 1.0000
14.750 1.6386 0.04132 0.03537 -0.0677 0.0135 1.0000
15.000 1.6390 0.04393 0.03804 -0.0671 0.0126 1.0000
15.250 1.6400 0.04658 0.04076 -0.0666 0.0119 1.0000
15.500 1.6404 0.04938 0.04364 -0.0663 0.0113 1.0000
15.750 1.6391 0.05250 0.04682 -0.0662 0.0106 1.0000
16.000 1.6368 0.05584 0.05024 -0.0663 0.0100 1.0000
16.250 1.6348 0.05924 0.05372 -0.0665 0.0096 1.0000
16.500 1.6327 0.06276 0.05732 -0.0669 0.0091 1.0000
16.750 1.6294 0.06655 0.06120 -0.0675 0.0087 1.0000
17.000 1.6251 0.07057 0.06531 -0.0683 0.0083 1.0000
17.250 1.6192 0.07491 0.06973 -0.0693 0.0079 1.0000
17.500 1.6135 0.07933 0.07425 -0.0704 0.0075 1.0000
17.750 1.6088 0.08369 0.07870 -0.0716 0.0072 1.0000
18.000 1.6032 0.08827 0.08337 -0.0730 0.0069 1.0000
18.250 1.5969 0.09302 0.08822 -0.0746 0.0066 1.0000
18.500 1.5896 0.09801 0.09329 -0.0764 0.0062 1.0000
18.750 1.5817 0.10318 0.09855 -0.0783 0.0059 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 562 AIRFOIL (e562-il)