AH-6-40-7 AIRFOIL (ah6407-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: AH-6-40-7 AIRFOIL (ah6407-il) Reynolds number: 500,000 Max Cl/Cd: 119.73 at α=1.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ah6407-il-500000.txt Download as CSV file: xf-ah6407-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: AH-6-40-7 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3389 0.10172 0.09956 -0.0260 1.0000 0.0170
-8.000 -0.3510 0.10060 0.09849 -0.0228 1.0000 0.0170
-7.750 -0.3631 0.09946 0.09739 -0.0197 0.9998 0.0170
-7.500 -0.3476 0.09360 0.09154 -0.0246 0.9973 0.0173
-7.250 -0.3305 0.09039 0.08833 -0.0251 0.9961 0.0178
-7.000 -0.3075 0.08656 0.08448 -0.0301 0.9939 0.0183
-6.750 -0.2854 0.08267 0.08059 -0.0356 0.9905 0.0189
-6.500 -0.2593 0.07840 0.07631 -0.0424 0.9874 0.0197
-6.250 -0.2269 0.07352 0.07140 -0.0515 0.9850 0.0211
-6.000 -0.1765 0.06705 0.06488 -0.0681 0.9812 0.0224
-5.750 -0.1387 0.06140 0.05917 -0.0780 0.9773 0.0225
-5.500 -0.1063 0.05395 0.05167 -0.0875 0.9749 0.0232
-5.250 -0.0746 0.05119 0.04887 -0.0919 0.9734 0.0241
-5.000 -0.0349 0.04785 0.04548 -0.0987 0.9722 0.0264
-4.750 0.0356 0.04052 0.03789 -0.1144 0.9715 0.0298
-4.500 0.0919 0.03171 0.02868 -0.1263 0.9707 0.0300
-4.250 0.1434 0.02069 0.01694 -0.1374 0.9697 0.0274
-4.000 0.1818 0.01758 0.01326 -0.1401 0.9647 0.0292
-3.750 0.2218 0.01433 0.00941 -0.1433 0.9604 0.0314
-3.500 0.2534 0.01376 0.00881 -0.1441 0.9541 0.0342
-3.250 0.2831 0.01293 0.00782 -0.1443 0.9465 0.0363
-3.000 0.3133 0.01213 0.00683 -0.1445 0.9408 0.0380
-2.750 0.3408 0.01197 0.00655 -0.1441 0.9325 0.0398
-2.500 0.3703 0.01072 0.00511 -0.1443 0.9257 0.0415
-2.250 0.3985 0.00997 0.00428 -0.1442 0.9169 0.0432
-2.000 0.4263 0.00960 0.00388 -0.1439 0.9078 0.0458
-1.750 0.4541 0.00927 0.00347 -0.1436 0.8985 0.0478
-1.500 0.4821 0.00898 0.00313 -0.1434 0.8879 0.0494
-1.250 0.5103 0.00875 0.00286 -0.1432 0.8775 0.0512
-1.000 0.5384 0.00858 0.00264 -0.1430 0.8672 0.0530
-0.750 0.5670 0.00832 0.00232 -0.1429 0.8558 0.0595
-0.500 0.5952 0.00819 0.00215 -0.1428 0.8434 0.0682
-0.250 0.6240 0.00786 0.00214 -0.1430 0.8306 0.1756
0.000 0.6520 0.00778 0.00213 -0.1429 0.8169 0.2282
0.500 0.7078 0.00754 0.00217 -0.1428 0.7844 0.3842
0.750 0.7285 0.00632 0.00219 -0.1412 0.7660 1.0000
1.000 0.7559 0.00645 0.00219 -0.1409 0.7451 1.0000
1.250 0.7832 0.00660 0.00220 -0.1406 0.7224 1.0000
1.500 0.8101 0.00678 0.00223 -0.1402 0.6967 1.0000
1.750 0.8369 0.00699 0.00229 -0.1399 0.6675 1.0000
2.000 0.8634 0.00724 0.00237 -0.1395 0.6341 1.0000
2.250 0.8894 0.00755 0.00247 -0.1390 0.5962 1.0000
2.500 0.9152 0.00790 0.00261 -0.1386 0.5550 1.0000
2.750 0.9409 0.00827 0.00279 -0.1382 0.5154 1.0000
3.250 0.9926 0.00900 0.00318 -0.1375 0.4502 1.0000
3.500 1.0187 0.00933 0.00338 -0.1372 0.4248 1.0000
3.750 1.0448 0.00965 0.00362 -0.1369 0.4036 1.0000
4.000 1.0710 0.00996 0.00385 -0.1366 0.3843 1.0000
4.250 1.0970 0.01027 0.00408 -0.1363 0.3671 1.0000
4.500 1.1231 0.01057 0.00433 -0.1360 0.3517 1.0000
4.750 1.1491 0.01087 0.00459 -0.1357 0.3372 1.0000
5.000 1.1747 0.01120 0.00485 -0.1353 0.3203 1.0000
5.250 1.2005 0.01149 0.00509 -0.1350 0.3017 1.0000
5.500 1.2261 0.01179 0.00535 -0.1347 0.2873 1.0000
5.750 1.2517 0.01209 0.00564 -0.1343 0.2723 1.0000
6.000 1.2771 0.01241 0.00592 -0.1340 0.2564 1.0000
6.250 1.3023 0.01273 0.00622 -0.1335 0.2412 1.0000
6.500 1.3266 0.01315 0.00656 -0.1330 0.2201 1.0000
6.750 1.3502 0.01366 0.00696 -0.1324 0.1883 1.0000
7.000 1.3658 0.01522 0.00791 -0.1308 0.0957 1.0000
7.250 1.3792 0.01710 0.00931 -0.1286 0.0281 1.0000
7.500 1.3999 0.01797 0.01019 -0.1273 0.0215 1.0000
7.750 1.4209 0.01872 0.01100 -0.1262 0.0188 1.0000
8.000 1.4382 0.01989 0.01230 -0.1244 0.0166 1.0000
8.250 1.4570 0.02079 0.01331 -0.1228 0.0158 1.0000
8.500 1.4749 0.02170 0.01433 -0.1212 0.0150 1.0000
8.750 1.4913 0.02270 0.01542 -0.1193 0.0142 1.0000
9.000 1.5065 0.02372 0.01653 -0.1173 0.0133 1.0000
9.250 1.5173 0.02506 0.01796 -0.1147 0.0124 1.0000
9.500 1.5181 0.02712 0.02016 -0.1104 0.0119 1.0000
9.750 1.5205 0.02891 0.02208 -0.1064 0.0116 1.0000
10.000 1.5283 0.03007 0.02335 -0.1033 0.0113 1.0000
10.250 1.5348 0.03142 0.02483 -0.1001 0.0111 1.0000
10.500 1.5402 0.03300 0.02653 -0.0970 0.0108 1.0000
10.750 1.5449 0.03474 0.02840 -0.0940 0.0106 1.0000
11.000 1.5492 0.03662 0.03041 -0.0912 0.0103 1.0000
11.250 1.5530 0.03866 0.03259 -0.0884 0.0101 1.0000
11.500 1.5562 0.04084 0.03493 -0.0858 0.0099 1.0000
11.750 1.5588 0.04306 0.03730 -0.0835 0.0097 1.0000
12.000 1.5606 0.04531 0.03967 -0.0813 0.0094 1.0000
12.250 1.5611 0.04772 0.04221 -0.0793 0.0092 1.0000
12.500 1.5607 0.05021 0.04481 -0.0776 0.0089 1.0000
12.750 1.5584 0.05313 0.04784 -0.0759 0.0087 1.0000
13.000 1.5536 0.05665 0.05150 -0.0743 0.0085 1.0000
13.250 1.5461 0.06069 0.05573 -0.0728 0.0084 1.0000
13.500 1.5366 0.06498 0.06022 -0.0717 0.0083 1.0000
13.750 1.5245 0.06975 0.06520 -0.0710 0.0083 1.0000
14.000 1.5111 0.07473 0.07039 -0.0710 0.0083 1.0000
14.250 1.4976 0.07990 0.07575 -0.0716 0.0083 1.0000
14.500 1.4824 0.08558 0.08163 -0.0729 0.0083 1.0000
14.750 1.4674 0.09152 0.08775 -0.0748 0.0083 1.0000
15.000 1.4530 0.09773 0.09415 -0.0774 0.0083 1.0000
15.250 1.4381 0.10444 0.10104 -0.0806 0.0083 1.0000
15.500 1.4234 0.11148 0.10825 -0.0844 0.0084 1.0000
15.750 1.4086 0.11888 0.11582 -0.0886 0.0084 1.0000
16.000 1.3933 0.12680 0.12392 -0.0935 0.0084 1.0000
16.250 1.3775 0.13547 0.13279 -0.0992 0.0086 1.0000
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Polar data table (+)
Polar graphs
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