AH 79-100 A AIRFOIL (ah79100a-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: AH 79-100 A AIRFOIL (ah79100a-il) Reynolds number: 200,000 Max Cl/Cd: 85.54 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ah79100a-il-200000.txt Download as CSV file: xf-ah79100a-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3837 0.09213 0.08872 -0.0446 1.0000 0.0562
-8.500 -0.3975 0.08954 0.08624 -0.0446 1.0000 0.0564
-8.250 -0.4181 0.08765 0.08445 -0.0426 1.0000 0.0565
-8.000 -0.4364 0.08533 0.08222 -0.0413 1.0000 0.0565
-7.750 -0.4377 0.08200 0.07892 -0.0356 1.0000 0.0580
-7.500 -0.4440 0.08045 0.07743 -0.0326 1.0000 0.0586
-7.250 -0.4501 0.07885 0.07587 -0.0301 1.0000 0.0595
-7.000 -0.4578 0.07674 0.07381 -0.0291 1.0000 0.0602
-6.750 -0.4471 0.07286 0.06994 -0.0332 0.9976 0.0623
-6.500 -0.4075 0.05298 0.04943 -0.0716 0.9877 0.0707
-6.250 -0.3776 0.03781 0.03341 -0.0815 0.9830 0.0429
-6.000 -0.3400 0.02820 0.02259 -0.0886 0.9804 0.0373
-5.750 -0.3056 0.02527 0.01920 -0.0911 0.9765 0.0380
-5.500 -0.2695 0.02389 0.01757 -0.0934 0.9721 0.0419
-5.250 -0.2293 0.02209 0.01528 -0.0959 0.9693 0.0444
-5.000 -0.1947 0.01993 0.01294 -0.0979 0.9659 0.0484
-4.750 -0.1619 0.01900 0.01189 -0.0991 0.9607 0.0525
-4.500 -0.1238 0.01811 0.01080 -0.1010 0.9574 0.0570
-4.250 -0.0852 0.01684 0.00958 -0.1035 0.9550 0.0637
-4.000 -0.0549 0.01621 0.00884 -0.1039 0.9491 0.0700
-3.750 -0.0188 0.01541 0.00807 -0.1057 0.9448 0.0832
-3.500 0.0211 0.01440 0.00721 -0.1085 0.9419 0.1138
-3.250 0.0629 0.01318 0.00665 -0.1120 0.9399 0.2572
-3.000 0.0899 0.01287 0.00664 -0.1120 0.9319 0.3570
-2.750 0.1296 0.01272 0.00657 -0.1142 0.9283 0.4178
-2.500 0.1704 0.01259 0.00641 -0.1165 0.9255 0.4563
-2.250 0.1974 0.01259 0.00642 -0.1161 0.9174 0.4844
-2.000 0.2336 0.01247 0.00632 -0.1174 0.9130 0.5121
-1.750 0.2676 0.01235 0.00620 -0.1183 0.9079 0.5350
-1.500 0.2965 0.01228 0.00614 -0.1181 0.9000 0.5579
-1.250 0.3321 0.01207 0.00592 -0.1191 0.8956 0.5784
-1.000 0.3578 0.01199 0.00587 -0.1184 0.8860 0.5949
-0.750 0.3906 0.01178 0.00566 -0.1188 0.8808 0.6128
-0.500 0.4163 0.01170 0.00561 -0.1181 0.8711 0.6289
-0.250 0.4477 0.01148 0.00540 -0.1182 0.8651 0.6472
0.000 0.4731 0.01138 0.00534 -0.1174 0.8544 0.6656
0.250 0.5002 0.01122 0.00522 -0.1168 0.8449 0.6838
0.500 0.5285 0.01102 0.00507 -0.1163 0.8363 0.7033
0.750 0.5539 0.01091 0.00503 -0.1155 0.8253 0.7267
1.000 0.5799 0.01075 0.00496 -0.1147 0.8153 0.7535
1.250 0.6061 0.01051 0.00481 -0.1137 0.8058 0.7887
1.500 0.6296 0.01022 0.00472 -0.1123 0.7932 0.8485
1.750 0.6650 0.00996 0.00455 -0.1135 0.7804 1.0000
2.000 0.6936 0.01001 0.00451 -0.1135 0.7677 1.0000
2.250 0.7218 0.01006 0.00451 -0.1134 0.7548 1.0000
2.500 0.7493 0.01012 0.00450 -0.1131 0.7408 1.0000
2.750 0.7765 0.01018 0.00450 -0.1127 0.7256 1.0000
3.000 0.8034 0.01027 0.00454 -0.1122 0.7101 1.0000
3.250 0.8302 0.01039 0.00463 -0.1118 0.6942 1.0000
3.500 0.8565 0.01051 0.00470 -0.1112 0.6767 1.0000
3.750 0.8826 0.01065 0.00474 -0.1105 0.6573 1.0000
4.000 0.9079 0.01082 0.00490 -0.1098 0.6362 1.0000
4.250 0.9335 0.01102 0.00505 -0.1092 0.6170 1.0000
4.500 0.9589 0.01126 0.00523 -0.1085 0.5983 1.0000
4.750 0.9837 0.01150 0.00545 -0.1078 0.5769 1.0000
5.000 1.0082 0.01179 0.00569 -0.1070 0.5549 1.0000
5.250 1.0321 0.01208 0.00595 -0.1061 0.5309 1.0000
5.500 1.0554 0.01241 0.00622 -0.1051 0.5055 1.0000
5.750 1.0782 0.01276 0.00651 -0.1041 0.4784 1.0000
6.000 1.1005 0.01313 0.00686 -0.1030 0.4497 1.0000
6.250 1.1221 0.01356 0.00722 -0.1018 0.4187 1.0000
6.500 1.1425 0.01406 0.00763 -0.1005 0.3811 1.0000
6.750 1.1602 0.01475 0.00811 -0.0988 0.3323 1.0000
7.000 1.1765 0.01561 0.00871 -0.0970 0.2771 1.0000
7.250 1.1918 0.01660 0.00942 -0.0952 0.2236 1.0000
7.500 1.2066 0.01769 0.01023 -0.0934 0.1744 1.0000
7.750 1.2208 0.01886 0.01115 -0.0916 0.1329 1.0000
8.000 1.2339 0.02014 0.01220 -0.0895 0.0958 1.0000
8.250 1.2421 0.02184 0.01363 -0.0868 0.0642 1.0000
8.500 1.2507 0.02342 0.01518 -0.0839 0.0485 1.0000
8.750 1.2558 0.02504 0.01677 -0.0807 0.0415 1.0000
9.000 1.2671 0.02617 0.01801 -0.0781 0.0374 1.0000
9.250 1.2755 0.02753 0.01942 -0.0755 0.0348 1.0000
9.500 1.2785 0.02974 0.02162 -0.0724 0.0327 1.0000
9.750 1.2911 0.03120 0.02320 -0.0704 0.0314 1.0000
10.000 1.3048 0.03281 0.02493 -0.0686 0.0302 1.0000
10.250 1.3185 0.03440 0.02665 -0.0669 0.0287 1.0000
10.500 1.3312 0.03597 0.02829 -0.0653 0.0273 1.0000
10.750 1.3446 0.03783 0.03019 -0.0640 0.0260 1.0000
11.250 1.3825 0.04405 0.03680 -0.0625 0.0246 1.0000
11.500 1.3902 0.04654 0.03957 -0.0605 0.0244 1.0000
11.750 1.3944 0.04935 0.04267 -0.0584 0.0242 1.0000
12.000 1.3940 0.05243 0.04605 -0.0561 0.0241 1.0000
12.250 1.3898 0.05586 0.04980 -0.0540 0.0241 1.0000
12.500 1.3819 0.05960 0.05383 -0.0520 0.0241 1.0000
12.750 1.3705 0.06363 0.05814 -0.0504 0.0242 1.0000
13.000 1.3563 0.06796 0.06274 -0.0493 0.0242 1.0000
13.250 1.3402 0.07270 0.06773 -0.0488 0.0243 1.0000
13.500 1.3224 0.07776 0.07303 -0.0490 0.0243 1.0000
13.750 1.3030 0.08348 0.07898 -0.0501 0.0244 1.0000
14.000 1.2826 0.08982 0.08553 -0.0519 0.0246 1.0000
14.250 1.2615 0.09687 0.09277 -0.0545 0.0247 1.0000
14.500 1.2433 0.10393 0.10002 -0.0575 0.0249 1.0000
14.750 1.2268 0.11024 0.10649 -0.0613 0.0251 1.0000
15.000 1.2092 0.11738 0.11382 -0.0662 0.0253 1.0000
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Polar data table (+)
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