AH-7-47-6 AIRFOIL (ah7476-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: AH-7-47-6 AIRFOIL (ah7476-il) Reynolds number: 500,000 Max Cl/Cd: 110.45 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ah7476-il-500000-n5.txt Download as CSV file: xf-ah7476-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AH-7-47-6 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.2131 0.09772 0.09532 -0.0588 0.9834 0.0086
-8.250 -0.2011 0.09412 0.09172 -0.0618 0.9811 0.0087
-8.000 -0.1875 0.09027 0.08787 -0.0654 0.9793 0.0088
-7.500 -0.1635 0.08397 0.08158 -0.0707 0.9721 0.0095
-7.250 -0.1422 0.08123 0.07884 -0.0747 0.9703 0.0103
-7.000 -0.1189 0.07739 0.07498 -0.0805 0.9689 0.0108
-6.750 -0.1045 0.07402 0.07162 -0.0838 0.9631 0.0110
-6.500 -0.0780 0.06955 0.06714 -0.0909 0.9608 0.0114
-6.250 -0.0434 0.06341 0.06096 -0.1019 0.9589 0.0121
-6.000 -0.0254 0.06177 0.05931 -0.1035 0.9528 0.0125
-5.750 0.0057 0.05873 0.05625 -0.1095 0.9501 0.0137
-5.500 0.0441 0.05394 0.05141 -0.1190 0.9473 0.0144
-5.250 0.0901 0.04732 0.04470 -0.1321 0.9409 0.0155
-5.000 0.1189 0.04582 0.04317 -0.1351 0.9383 0.0163
-4.750 0.1646 0.04189 0.03915 -0.1442 0.9365 0.0179
-4.500 0.2366 0.03320 0.03019 -0.1621 0.9323 0.0195
-4.250 0.2610 0.03277 0.02975 -0.1628 0.9278 0.0203
-4.000 0.3136 0.02872 0.02549 -0.1713 0.9255 0.0231
-3.750 0.3640 0.02424 0.02074 -0.1791 0.9226 0.0228
-3.250 0.4474 0.01885 0.01482 -0.1880 0.9128 0.0259
-3.000 0.4883 0.01644 0.01204 -0.1916 0.9059 0.0261
-2.750 0.5267 0.01464 0.00989 -0.1942 0.8983 0.0260
-2.500 0.5611 0.01347 0.00845 -0.1958 0.8900 0.0261
-2.000 0.6257 0.01201 0.00659 -0.1979 0.8717 0.0272
-1.750 0.6574 0.01130 0.00568 -0.1987 0.8623 0.0265
-1.500 0.6882 0.01075 0.00497 -0.1993 0.8533 0.0261
-1.250 0.7181 0.01033 0.00442 -0.1997 0.8447 0.0261
-0.750 0.7768 0.00971 0.00365 -0.2003 0.8281 0.0267
-0.500 0.8058 0.00950 0.00338 -0.2005 0.8202 0.0272
-0.250 0.8345 0.00933 0.00318 -0.2007 0.8115 0.0276
0.000 0.8631 0.00920 0.00301 -0.2008 0.8025 0.0280
0.250 0.8911 0.00915 0.00293 -0.2008 0.7920 0.0288
0.500 0.9188 0.00913 0.00290 -0.2007 0.7810 0.0295
0.750 0.9466 0.00905 0.00279 -0.2007 0.7652 0.0295
1.000 0.9732 0.00905 0.00272 -0.2003 0.7384 0.0295
1.250 0.9973 0.00923 0.00268 -0.1994 0.6871 0.0295
1.500 1.0185 0.00967 0.00277 -0.1979 0.6208 0.0296
1.750 1.0402 0.01016 0.00296 -0.1966 0.5606 0.0297
2.000 1.0629 0.01058 0.00311 -0.1956 0.5083 0.0302
2.500 1.1111 0.01126 0.00343 -0.1942 0.4440 0.0323
2.750 1.1358 0.01155 0.00362 -0.1936 0.4227 0.0339
3.000 1.1610 0.01178 0.00379 -0.1932 0.4058 0.0360
3.250 1.1866 0.01197 0.00397 -0.1927 0.3936 0.0411
3.500 1.2132 0.01206 0.00423 -0.1926 0.3819 0.1503
4.000 1.2613 0.01142 0.00490 -0.1915 0.3503 1.0000
4.250 1.2854 0.01176 0.00514 -0.1908 0.3259 1.0000
4.500 1.3061 0.01243 0.00551 -0.1896 0.2703 1.0000
4.750 1.3245 0.01334 0.00605 -0.1880 0.2077 1.0000
5.000 1.3452 0.01404 0.00653 -0.1867 0.1700 1.0000
5.250 1.3670 0.01458 0.00694 -0.1857 0.1453 1.0000
5.500 1.3882 0.01518 0.00741 -0.1845 0.1175 1.0000
5.750 1.4082 0.01590 0.00793 -0.1832 0.0836 1.0000
6.000 1.4279 0.01661 0.00850 -0.1818 0.0609 1.0000
6.250 1.4494 0.01711 0.00898 -0.1806 0.0508 1.0000
6.500 1.4697 0.01771 0.00955 -0.1793 0.0376 1.0000
6.750 1.4868 0.01861 0.01029 -0.1774 0.0172 1.0000
7.000 1.5058 0.01928 0.01098 -0.1757 0.0136 1.0000
7.250 1.5252 0.01986 0.01163 -0.1741 0.0118 1.0000
7.500 1.5443 0.02043 0.01228 -0.1725 0.0109 1.0000
7.750 1.5621 0.02107 0.01301 -0.1706 0.0100 1.0000
8.000 1.5786 0.02179 0.01381 -0.1686 0.0092 1.0000
8.250 1.5921 0.02273 0.01485 -0.1660 0.0083 1.0000
8.500 1.6069 0.02348 0.01567 -0.1636 0.0078 1.0000
8.750 1.6202 0.02419 0.01648 -0.1610 0.0074 1.0000
9.000 1.6323 0.02504 0.01741 -0.1582 0.0070 1.0000
9.250 1.6438 0.02594 0.01842 -0.1554 0.0066 1.0000
9.500 1.6544 0.02693 0.01949 -0.1525 0.0063 1.0000
9.750 1.6643 0.02798 0.02063 -0.1496 0.0060 1.0000
10.000 1.6723 0.02918 0.02193 -0.1465 0.0058 1.0000
10.250 1.6760 0.03075 0.02360 -0.1429 0.0056 1.0000
10.500 1.6780 0.03248 0.02546 -0.1391 0.0054 1.0000
10.750 1.6839 0.03393 0.02704 -0.1360 0.0053 1.0000
11.000 1.6896 0.03544 0.02869 -0.1331 0.0051 1.0000
11.250 1.6940 0.03708 0.03046 -0.1301 0.0050 1.0000
11.500 1.6977 0.03884 0.03235 -0.1273 0.0048 1.0000
11.750 1.7011 0.04066 0.03431 -0.1246 0.0046 1.0000
12.000 1.7040 0.04258 0.03636 -0.1220 0.0044 1.0000
12.250 1.7061 0.04464 0.03855 -0.1196 0.0043 1.0000
12.500 1.7073 0.04686 0.04089 -0.1173 0.0041 1.0000
12.750 1.7068 0.04934 0.04351 -0.1151 0.0040 1.0000
13.000 1.7050 0.05202 0.04633 -0.1130 0.0040 1.0000
13.250 1.7019 0.05495 0.04940 -0.1112 0.0039 1.0000
13.500 1.6980 0.05810 0.05269 -0.1096 0.0038 1.0000
13.750 1.6923 0.06163 0.05637 -0.1082 0.0038 1.0000
14.000 1.6851 0.06550 0.06038 -0.1072 0.0037 1.0000
14.250 1.6761 0.06983 0.06487 -0.1065 0.0036 1.0000
14.500 1.6657 0.07461 0.06981 -0.1064 0.0036 1.0000
14.750 1.6532 0.07994 0.07531 -0.1066 0.0035 1.0000
15.000 1.6385 0.08585 0.08141 -0.1074 0.0035 1.0000
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Polar data table (+)
Polar graphs
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