AH 79-100 A AIRFOIL (ah79100a-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: AH 79-100 A AIRFOIL (ah79100a-il) Reynolds number: 100,000 Max Cl/Cd: 61.95 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ah79100a-il-100000.txt Download as CSV file: xf-ah79100a-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: AH 79-100 A AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3531 0.10222 0.09722 -0.0359 1.0000 0.1056
-8.750 -0.3660 0.10025 0.09536 -0.0380 1.0000 0.1103
-8.500 -0.3945 0.09939 0.09469 -0.0404 1.0000 0.1114
-8.250 -0.3612 0.09401 0.08924 -0.0362 1.0000 0.1159
-8.000 -0.3624 0.09199 0.08728 -0.0351 1.0000 0.1217
-7.750 -0.3839 0.09068 0.08613 -0.0344 1.0000 0.1243
-7.500 -0.4105 0.08977 0.08538 -0.0324 1.0000 0.1251
-7.250 -0.4427 0.08754 0.08329 -0.0375 1.0000 0.1260
-7.000 -0.4085 0.08482 0.08052 -0.0263 1.0000 0.1328
-6.750 -0.4251 0.08305 0.07886 -0.0261 1.0000 0.1379
-6.500 -0.4512 0.07822 0.07411 -0.0375 1.0000 0.1412
-6.250 -0.4387 0.07757 0.07351 -0.0256 1.0000 0.1448
-6.000 -0.4475 0.07207 0.06799 -0.0380 1.0000 0.1562
-5.750 -0.4421 0.07062 0.06662 -0.0307 1.0000 0.1582
-5.500 -0.4017 0.04333 0.03776 -0.0657 1.0000 0.0846
-5.250 -0.3986 0.04710 0.04219 -0.0599 1.0000 0.1144
-5.000 -0.3432 0.03253 0.02574 -0.0717 1.0000 0.0731
-4.750 -0.3167 0.03026 0.02312 -0.0727 1.0000 0.0763
-4.500 -0.2872 0.02763 0.01991 -0.0737 1.0000 0.0774
-4.250 -0.2596 0.02601 0.01778 -0.0741 1.0000 0.0816
-4.000 -0.2333 0.02423 0.01577 -0.0744 1.0000 0.0862
-3.750 -0.2084 0.02309 0.01451 -0.0742 1.0000 0.0911
-3.500 -0.1836 0.02213 0.01341 -0.0740 1.0000 0.0997
-3.250 -0.1538 0.02132 0.01253 -0.0747 0.9983 0.1105
-3.000 -0.1128 0.02037 0.01178 -0.0776 0.9932 0.1334
-2.750 -0.0701 0.01932 0.01112 -0.0810 0.9884 0.2006
-2.500 -0.0309 0.01871 0.01146 -0.0838 0.9817 0.4127
-2.250 0.0060 0.01918 0.01196 -0.0856 0.9734 0.4823
-2.000 0.0443 0.01967 0.01244 -0.0875 0.9657 0.5324
-1.750 0.0794 0.01997 0.01276 -0.0887 0.9570 0.5700
-1.500 0.1130 0.02019 0.01297 -0.0898 0.9481 0.6046
-1.250 0.1526 0.02033 0.01312 -0.0918 0.9408 0.6384
-1.000 0.1829 0.02036 0.01315 -0.0923 0.9306 0.6652
-0.750 0.2226 0.02032 0.01314 -0.0943 0.9234 0.6927
-0.500 0.2553 0.02023 0.01307 -0.0951 0.9133 0.7181
-0.250 0.2887 0.02015 0.01300 -0.0960 0.9037 0.7458
0.000 0.3300 0.01990 0.01280 -0.0982 0.8966 0.7772
0.250 0.3593 0.01974 0.01272 -0.0983 0.8860 0.8110
0.500 0.4041 0.01919 0.01233 -0.1007 0.8808 0.8650
0.750 0.4486 0.01891 0.01213 -0.1043 0.8698 1.0000
1.000 0.4966 0.01880 0.01189 -0.1084 0.8619 1.0000
1.250 0.5393 0.01865 0.01164 -0.1111 0.8529 1.0000
1.500 0.5734 0.01866 0.01158 -0.1121 0.8418 1.0000
1.750 0.6180 0.01832 0.01120 -0.1145 0.8357 1.0000
2.000 0.6474 0.01836 0.01121 -0.1145 0.8232 1.0000
2.250 0.6783 0.01831 0.01114 -0.1146 0.8111 1.0000
2.500 0.7110 0.01815 0.01100 -0.1148 0.8000 1.0000
2.750 0.7478 0.01779 0.01062 -0.1154 0.7910 1.0000
3.000 0.7759 0.01777 0.01062 -0.1149 0.7776 1.0000
3.250 0.8045 0.01770 0.01059 -0.1143 0.7640 1.0000
3.500 0.8335 0.01759 0.01050 -0.1138 0.7499 1.0000
3.750 0.8624 0.01747 0.01040 -0.1132 0.7354 1.0000
4.000 0.8910 0.01739 0.01034 -0.1125 0.7204 1.0000
4.250 0.9197 0.01725 0.01023 -0.1118 0.7041 1.0000
4.500 0.9482 0.01709 0.01006 -0.1109 0.6864 1.0000
4.750 0.9728 0.01713 0.01014 -0.1096 0.6649 1.0000
5.000 0.9999 0.01711 0.01013 -0.1087 0.6448 1.0000
5.250 1.0253 0.01723 0.01027 -0.1076 0.6234 1.0000
5.500 1.0505 0.01737 0.01041 -0.1065 0.6006 1.0000
5.750 1.0742 0.01757 0.01062 -0.1051 0.5754 1.0000
6.000 1.0975 0.01779 0.01085 -0.1037 0.5484 1.0000
6.250 1.1194 0.01807 0.01111 -0.1021 0.5187 1.0000
6.500 1.1398 0.01841 0.01144 -0.1004 0.4859 1.0000
6.750 1.1591 0.01880 0.01178 -0.0985 0.4504 1.0000
7.000 1.1764 0.01931 0.01226 -0.0964 0.4095 1.0000
7.250 1.1911 0.02000 0.01280 -0.0940 0.3624 1.0000
7.500 1.2024 0.02097 0.01352 -0.0912 0.3059 1.0000
7.750 1.2106 0.02229 0.01448 -0.0883 0.2457 1.0000
8.000 1.2172 0.02394 0.01573 -0.0854 0.1905 1.0000
8.250 1.2220 0.02587 0.01729 -0.0824 0.1461 1.0000
8.500 1.2255 0.02804 0.01915 -0.0791 0.1128 1.0000
8.750 1.2308 0.03012 0.02109 -0.0761 0.0905 1.0000
9.000 1.2381 0.03224 0.02304 -0.0734 0.0764 1.0000
9.250 1.2525 0.03409 0.02497 -0.0714 0.0671 1.0000
9.500 1.2768 0.03676 0.02763 -0.0709 0.0608 1.0000
9.750 1.2970 0.03878 0.02976 -0.0699 0.0558 1.0000
10.000 1.3241 0.04199 0.03296 -0.0702 0.0514 1.0000
10.250 1.3440 0.04471 0.03608 -0.0690 0.0495 1.0000
10.500 1.3607 0.04784 0.03958 -0.0676 0.0482 1.0000
10.750 1.3715 0.05121 0.04334 -0.0656 0.0474 1.0000
11.000 1.3759 0.05478 0.04732 -0.0632 0.0471 1.0000
11.250 1.3728 0.05831 0.05124 -0.0602 0.0469 1.0000
11.500 1.3638 0.06173 0.05501 -0.0569 0.0468 1.0000
11.750 1.3507 0.06529 0.05890 -0.0538 0.0467 1.0000
12.000 1.3345 0.06910 0.06301 -0.0513 0.0466 1.0000
12.250 1.3145 0.07346 0.06767 -0.0495 0.0468 1.0000
12.500 1.2916 0.07839 0.07289 -0.0487 0.0472 1.0000
12.750 1.2665 0.08397 0.07874 -0.0490 0.0477 1.0000
13.000 1.2397 0.09034 0.08534 -0.0506 0.0483 1.0000
13.250 1.2123 0.09748 0.09270 -0.0534 0.0491 1.0000
13.500 1.1858 0.10540 0.10078 -0.0574 0.0499 1.0000
13.750 1.1605 0.11399 0.10950 -0.0623 0.0507 1.0000
14.000 1.1395 0.12284 0.11844 -0.0674 0.0515 1.0000
14.250 1.1277 0.13098 0.12660 -0.0710 0.0523 1.0000
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Polar data table (+)
Polar graphs
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